Georgia Institute of Technology Mile High Yellow Jackets

Transcription

1 Project A.P.E.S. Preliminary Design Review (PDR) Documentation

2 Table of Contents 1. Team Overview Team Summary Mission Statement Launch Vehicle Summary Payload Summary Changes since PDR Changes to the Launch Vehicle System Changes to the Payload and Flight Systems Design Changes to the Activity Plan Launch Vehicle Overview System Design Overview Recovery System Booster Section imps Integrated Modular Payload System Vespula Mass Breakdown Vespula Overall Dimension Launch Vehicle Performance Predictions Flight Simulation AeroTech L1390G-P Simulated Thrust Curve Stability Margin Drift Profile Simulation Kinetic Energy Upon Landing Launch Vehicle Testing Subscale Testing Launch System and Platform Payload Introduction to the Experiment and Payload Concept Features & Definition Overview of the Experiment of 94

3 8. Flight Systems Flight Avionics Power Systems Telemetry and Recovery Integration Sensing Capabilities De-Scope Options Safety General Safety Payload Hazards Vehicle Safety Budget Project Scheduling Educational Outreach References Appendix 1: Project A.P.E.S. Schedule Appendix 2: Launch Checklist Appendix 3: Ground Test Plan Appendix 4: Derivation of Force Equations for a Ferrite Platform Appendix 5: Flight Computer Schematic Appendix 6: Material Safety Data Sheets (MSDS) of 94

4 1. Team Overview 1.1. Team Summary Team Summary School Name Team Name Project Title Active Platform Electromagnetic Stabilization (A.P.E.S.) Launch vehicle Name Vespula Project Lead Richard Safety Officer Matt Team Advisors Dr. Eric Feron, Dr. Marilyn Wolf NAR Section Primary: Southern Area Launch vehiclery (SoAR) #571 Secondary: GA Tech Ramblin Launch vehicle Club #701 NAR Contact Primary: Matthew Vildzius Secondary: Jorge Blanco 1.2. Mission Statement To maintain a sustainable team dedicated to the gaining of knowledge through the designing, building, and launching of reusable launch vehicles with innovative payloads in accordance with the NASA University Student Launch Initiative Guidelines Launch Vehicle Summary The Vespula launch vehicle weighs approximately 40 pounds and the preferred motor is the AeroTech L1390G-P. The structure of the launch vehicle features a rib-and-stringer design covered by a thin skin to minimize weight. The recovery system utilizes a 36 drogue parachute slowing the launch vehicle down to 120 feet per second (fps) and a 120 main parachute to slow the launch vehicle down to 25 fps Payload Summary The will design, build, test, and fly an electromagnetically levitated plate within their launch vehicle. This plate will be stabilized against the motion of the launch vehicle providing a vibration-free environment for a theoretical payload in an experiment known as A.P.E.S., or Active Platform Electromagnetic Stabilization. Flight Systems will utilize an ATmega 2560 for all data collection activities and the ARM Cortex M3 for the A.P.E.S. control law implementation. 4 of 94

5 2. Changes since PDR 2.1. Changes to the Launch Vehicle System No changes were made to the launch vehicle 2.2. Changes to the Payload and Flight Systems Design The following changes have been made since the Proposal: Removal of the accelerometers from the feedback loop for the control system The laser Infrared systems have been substituted with a Camera Cube and IR Sensor The design paradigm of the platform has changed The Universal Mounting System is now known as the Universal Mounting Bracket Changed the Flight Computer from the Parallax Propeller Chip to the ATmega Changes to the Activity Plan No changes were made to the activity plan. 5 of 94

6 3. Launch Vehicle 3.1. Overview The purpose of the launch vehicle is to carry a payload up one mile in altitude and safely return it to the surface of the Earth. Additionally, the launch vehicle will also be designed to carry a wide range of possible experiments so that the launch vehicle can be reused in the future. The overall design is to be as flexible as possible, encouraging reuse for future research and multiple launches. The chosen launch vehicle design would provide an adequate amount of space for a variety of payload designs. Prior to test flights, extensive ground testing will be performed to verify successful integration of the payload into the fully assembled launch vehicle. A subscale test flight will occur in December 2011 to test the launch vehicle s skin design and a full scale test will be performed in the spring. The objective of the full scale test launch, from a vehicle perspective, is to verify the recovery system with delayed apogee ejection and the integrity of the overall structure during flight. These tests will ensure a successful launch and recovery of the vehicle to meet all requirements for the final launch Though a kit launch vehicle would be easier to construct, a custom internal structure was designed to have lower mass and lower cost. The launch vehicle will have a 5 inch outer diameter to fit the chosen nose cone. The vehicle's diameter of 5 inches supports a wide range of payloads on a level two motor, while providing support for up to a 40-inch-long payload bay. Strength, durability, and safety are ensured as the structure is solely composed of fiberglass components. Mission Success Criteria The criteria for mission success are shown in Table 1. 6 of 94

7 Table 1: Mission Success Criteria Requirement Provide a suitable environment for the payload To fly as close to a mile in altitude as possible without exceeding 5,600 ft. The vehicle must be reusable Design feature to satisfy that requirement The payload requires a steady, but randomly vibrating platform to test the APES system. Unsteadiness in the motor's thrust and launch vehicle aerodynamics cause vibrations. A motor will be chosen to propel the vehicle to a mile in altitude The structure will be robust enough to handle any loading encountered during the flight Requirement Verification By measuring the acceleration with the payload's accelerometers Through the use of barometric altimeters Through finite element analyses and structural ground testing of components Success Criteria The APES system dampens out a recordable amount of vibration. The altimeters record an altitude less than 5,600 ft The vehicle survives the flight with no damage 3.2. System Design Overview Table 2, Table 3, and Table 4list the derived system-level requirements in order to meet the success criteria. 7 of 94

8 Table 2: System design requirement overview Requirement The launch vehicle shall carry a science or engineering payload The launch vehicle shall deliver the payload to an altitude of 1 mile above ground level The launch vehicle shall carry one approved altimeter for recording purposes The recovery system shall meet all requirements listed in requirement 4 in the handbook The recovery system electronics shall be shielded from all interference The launch vehicle shall have a pad stay time of one (1) hour The launch vehicle shall have aerodynamic stability before leaving the launch rail The launch vehicle shall remain subsonic throughout flight The launch vehicle shall be recoverable and reusable The launch vehicle shall stage the deployment of its recovery devices Design feature that will satisfy that requirement The launch vehicle will carry the A.P.E.S. experiment in the integrated modular payload section (imps) The motor will be chosen per the final launch vehicle mass An approved altimeter will be included in the recovery design The recovery system will feature a drogue and main parachute Faraday shielding will be incorporated into design to protect electronics from payload interference The hardware and battery will be able to function after remaining on the pad for a hour The launch vehicle will have launch buttons mounted to the booster section and will launch from a rail of adequate length The motor will be chosen per the final launch vehicle mass The recovery system will feature a drogue and main parachute The recovery system will feature a drogue and main parachute Requirement Verification The A.P.E.S. experiment will undergo extensive ground testing prior to flight testing Verification via OpenRocket vehicle simulations of the design Engineering inspection from manufacturer Ground testing of the independent recovery systems and flight testing of the integrated system Ground testing of the independent recovery systems and flight testing of the integrated system Ground testing of the integrated system OpenRocket vehicle Simulations Through OpenRocket vehicle simulations of the design Ground testing of the independent recovery systems and flight testing of the integrated system Ground testing of the independent recovery systems and flight testing of the integrated system 8 of 94

9 Table 3: System design requirement overview Requirement Removable shear pins shall be used for both the main and drogue chute compartments The launch vehicle shall have a maximum of four (4) independent or tethered sections Each section shall have a maximum kinetic energy of 75 ftlb f All sections shall be designed to recover within 2,500 feet of the launch pad assuming 15 mph wind The launch vehicle shall be capable of being prepared for flight at the launch site within 2 hours from the time the waiver opens The launch vehicle shall be launched from a standard firing system using a 10 second countdown The launch vehicle shall require no external circuitry or special ground support equipment to initiate the launch other than what is provided by the range Design feature that will satisfy that requirement Plastic shear pins are designed to be installed in the recovery compartments There are three (3) sections: nosecone, payload, and booster The recovery system will feature a drogue and main parachute The recovery system will feature a drogue and main parachute Assembly Checklist NAR launch regulations will be followed The vehicle will implement standard ground support equipment Requirement Verification Ground testing of the independent recovery systems and flight testing of the integrated system Engineering inspection Analysis and simulation OpenRocket vehicle simulations Construction testing and instructions will be made before launch day Range Safety Officer Analysis and Ground Testing 9 of 94

10 Table 4: System design requirement overview Requirement The launch vehicle shall use a commercially available solid motor The total impulse provided by the launch vehicle shall not exceed 5,120 N-s Design feature that will satisfy that requirement An L-class NAR approved motor will be incorporated into the design An L-class NAR approved motor will be incorporated into the design Requirement Verification Engineering inspection from manufacturer Engineering inspection from manufacturer 3.3. Recovery System The recovery system is intended to mitigate difficulties encountered due to variable wind speeds and to prevent destruction from impact. These objectives will be accomplished with a dualdeployment system. OpenRocket vehicle analysis of the drogue chute indicates that the maximum descent rate will be 120 fps with deployment at apogee. Main chute deployment will slow the launch vehicle to a maximum descent rate of 25 fps and will be deployed at 650 ft AGL altitude. The drogue chute is assumed to have a C D of 1.2, and the main chute is assumed to have a C D of 1.4. Figure 1 shows a section drawing of the launch vehicle and where the parachutes are contained. Figure 1: Section View of Launch vehicle Parachute Dimensions The drogue chute will be placed inside the nose cone which has an outer diameter of 5 inches, an inner diameter of 4.7 inches, and a length of 25.5 inches. A bulkhead connected only to shock cord will serve to take the impulse of the gun powder blast and prevent the drogue chute from being stuck inside of the nose cone. Once ejected, the inertia of the nose cone will allow the 10 of 94

11 drogue chute to be pulled freely from storage. The main chute will be housed in a cylindrical compartment attached to the thrust plate within the launch vehicle. The main chute compartment has a 5 inch outer diameter and a height of 15 inches. Shock cords will connect the parachutes to all sections of the launch vehicle such that in the recovery phase the entire system remains a single unit. For the main chute, the shock cord is attached to a U-Bolt that is on the reverse side of the thrust plate, and will be connected through a steel wire drilled into a fiberglass bulkhead. Similarly, the shock cord of the drogue parachute is connected to the nose cone and to the steel wire on the upper end of the payload section bulkhead. Table 5 outlines the dimensions and the weights of the parachutes. Table 5: Parachute parameters Main Parachute Drogue Parachute Dimensions 10 ft diameter 3 ft diameter Surface Area ft ft 2 Estimated C D Weight 2.3 lb 1.5 lb Target Descent Rate 25 fps 120 fps Ejection Charges Black powder masses were calculated using Equation ( 1 ) with variables defined in Table 6.. W = VDP RT ( 1 ) 11 of 94

12 Table 6: Ejection charge equation variables Variable Description Units W Weight of the black powder in pound mass 454 Wgram V Volume of the container to be pressurized in 3 DP Pressure Differential psia R Gas Combustion Constant for black powder ft lb lb R T Gas Combustion Temperature 3307 R Volume, V, is set by the design, while the black powder sets the gas constant and temperature. In order to find the pressurization, the strength and number of shear pins that will hold the parachute compartments together is needed. A quarter-inch shear pin can take up to 35 pounds of shear force before it fails. The two compartments will be held together with four shear pins, thus implying only a force of 70 pounds per compartment is needed to achieve separation. However, accounting for frictional resistance from the tubes, 87.5 pounds of force per compartment will be targeted. The corresponding amounts of black powder are summarized in Table 7. Figure 2 shows how the two (2) chutes will be detach throughout the launch vehicle. m f Table 7: Ejection pressurization and black powder charge Main Parachute Drogue Parachute Total Pressurization 24.7 psia 23.7 psia Differential Pressurization 10.0 psia 9.00 psia Amount of Black Powder 1.52 grams 0.91 grams Figure 2: Parachutes after Deployment Full Altimeters The StratoLogger collects flight data at a rate of twenty samples per second throughout the flight and stores them for later download to a computer. The altimeter is capable of recording flights of 12 of 94

13 up to 100,000 feet altitude. A picture of the StratoLogger is shown in Figure 3. Two (2) altimeters will be used, each with an independent power supply. The system for each altimeter will be set up as shown in Figure 4, while the pin connections to be used are shown in Table 8. Figure 3: Electrical schematics for recovery system 0 13 of 94

14 Figure 4: Electrical schematic of Stratologger Table 8: Summary of the electrical schematics for recovery system Port Name Description A Battery Port Connect to a 9 V power source B Power Switch Port Connect to a power switch C Main Chute Port Connect to Main Chute E-matches D Drogue Chute Port Connect to Drogue Chute E-matches G Beeper Audibly reports settings, status, etc. via a sequence of beeps Assembly Launch Vehicle Recovery Both the main and drogue parachutes are legacy hardware and the casings are made of G10 fiberglass. The bulkhead that goes underneath the main parachute will be made out of aluminum, since it will serve a dual function as the thrust plate. Four shear pins will hold the assembly in launch configuration. The ejection charge will be directed with the use of PVC end-caps to protect the casings from thermal shock. Shredded cellulose will be used to protect the chutes 14 of 94

15 from thermal shock. The charges will be ignited by an e-match. Each casing will undergo two tests: a distance test and a feasibility test that will estimate the mass to be placed both forward and aft of each compartment. These tests will confirm whether the black powder charges are enough to achieve separation. Supporting equipment will be provided to prevent launch vehicle parts or debris from hitting anyone. Recovery Ground Test Assembly will be straightforward. The main coupler will be fixed to the compartment using epoxy. The top half of the compartment will then slide onto the coupler up to the inner wall of the compartment and will be held in place by shear pins. The ejection charges will be placed on the outer side of the payload bulkheads and will be detonated by an e-match, which will be manually controlled by a switch box at a safe distance from the test structure. The actual chutes will be inside the compartments; a NOMEX/cloth shield will protect them. Testing For testing to be considered a success, it must meet all of the success criteria (shown in Table 10. Table 9). Meeting some of the criteria can be considered partial success. Only if none of the criteria are met, or if one of the failure modes occurs, would the test be considered a failure. The failure modes are shown in Table 10. Table 9: Success Criteria Success Criteria Ejection charge ignites Shear pins break Launch vehicle moves half the distance of shock cord 15 of 94

16 Table 10: Failure Modes Failure Criteria The fiberglass or the tube coupler shatters due to the charge The shear pins don t shear, and the launch vehicle stays intact The NOMEX/Cloth shield fails and the parachute is burned Ematches fail to ignite black powder 3.4. Booster Section The booster system of the launch vehicle uses traditional motors combined with a unique structure. The main focus is to have a highly integrated design and to lighten the total weight of the launch vehicle. The primary components of launch vehicle booster section would be the launch vehicle motor itself and the thrust retention system. Both of these primary components are shown in Figure 5 and Figure 6 below. Figure 5: AeroTech 75/3840 Motor Casing Figure 6: Thrust Retention System The Motor Retention System (MRS) can be broken down to several modular parts. The first part is the thrust plate located at the top of the MRS. The thrust plate will provide contact area for the ignited launch vehicle motor to provide thrust to the rest of the launch vehicle. The thrust plate will also be part of the recovery system for chute deployment as it will serve as a bulkhead with a U-bolt for attachment of the main chute. Finally, thrust plate will prevent the motor from penetrating through the booster section and it will provide torsional rigidity to the mounting rods that run along the MRS. The assembly of the thrust plate with the motor casing and mounting rods is shown in Figure of 94

17 Figure 7: Thrust Plate Assembly Two (2) major components of the MRS are the support plates at the rear of the booster section. Their main purpose is to prevent the motor from falling out of the launch vehicle. The secondary function is to provide support for three (3) aluminum (or plastic) U-Channel tabs. These tabs are slots that will be used to attach the fins to the booster section. The last two (2) major components are the centering rings that provide additional support to the fin tabs and recovery section tubes. The rear support assembly and the total assembly including the recovery system are shown in Figure 8 and Figure 9. Figure 8: Rear Support Assembly Figure 9: Booster Section with Recovery 17 of 94

18 Material Requirements Most of the booster section will be constructed from plywood and sheets of aluminum. These materials will reduce the weight of the system while maintaining structural integrity. However, the thrust plate will require a stronger aluminum alloy to minimize deformation. Manufacturing and Quality Assurance Quality will be assured by utilizing proper facilities that are appropriate for the materials and part designs of the booster section. First, the thrust plate and rear support plates require complex cutting. This will be done using a CNC milling machine provided by the Georgia Tech School of Mechanical Engineering Invention Studio. Also, a water jet will be used for cutting out the plywood for the centering rings. Afterwards, the booster section will be assembled using epoxy and hex nuts on threaded rod. For motor preparation, the rear support plates will be removed by unscrewing the hex nuts. Afterwards, the motor propellant will be placed inside the motor casing and the rear support plates will be reattached. The booster section will undergo several test firings to ensure reliability and quality for the official launch date. Five motors will be used for testing and the final launch. Furthermore, the MRS will undergo torsional buckling tests for structural integrity. Finite Element Analysis Finite Element Analysis (or FEA) is a method of finding approximate solutions to partial differential equations 0. This method is helpful in analyzing complex structures for design and development analysis. Solidworks contains a design validation tool that carries out basic FEA, which was used to structurally analyze the thrust plate and the overall booster section. The force applied is 408 lb f, which is the maximum thrust expected for an Aerotech L1390G-P. The first simulation is with only the thrust plate shown in Figure 10 through Figure 13. It resulted in a maximum displacement of inches and an induced maximum stress of 10,053 psi. The next simulation is for the entire booster section (see Figure 14 andfigure 15). It is important to note both figures have deformation scaled to a high level so that it appears as if the rods are buckling when they are not. To reduce complexity, the shape was generalized as a single part similar to the actual booster section which included the thrust plate, mounting rods, fin centering 18 of 94

19 ring, and the two support plates. The material used for the simulation was Aluminum 6061-T6. This simulation resulted in a maximum displacement of inches and an induced maximum stress of 9,620 psi. The results of these simulations are summarized in Table 11 below. Table 11: FEA results for the thrust plate and the assembly Force Applied (lb f ) Max. Displacement (inches) Max. Stress (psi) Factor of Safety Thrust Plate , Assembly , Figure 10: Thrust Plate Displacement (top-view) Figure 11: Thrust Plate Displacement (bottom-view) Figure 12: Thrust Plate Stresses (top-view) Figure 13: Thrust Plate Stress (bottom-view) 19 of 94

20 Figure 14: Booster section stresses Figure 15: Booster section displacement 3.5. Fin Overview Fins are used to keep the launch vehicle stable and the flight path straight. The fins will be made out of a composite honeycomb core with carbon fiber skins on both sides. This sandwich design provides strength to the fin structure while reducing weight. The fins will connect to the launch vehicle structure through plastic U-channels attached to the centering rings. The U-channels will run across the entire length of the fin providing the maximum contact area and epoxy will be used to attach the U-channel and the fin. During flight, if the drag becomes too great the fins can detach from the launch vehicle structure due to the high moment acting at the interface between the fins and the structure of the launch vehicle. Calculations of the max drag force per fin was performed utilizing Equation ( 2 ) and the moment was calculated as the product of force and the distance from the tip of the fin. D = 1 2 rv2 C D A ( 2 ) Due to the small or zero angle of attack, compressibility effects are negligible and therefore neglected 0. Further, the maximum velocity occurs at an altitude of 1,015 feet, which corresponds to a density of lb m /ft 3. The results are summarized in Table of 94

21 Table 12: Drag calculation values Variable Value Density lb m /ft 3 Velocity ft/sec Cross section area ft 2 C D Max drag force per fin Moment 8.9 lb f 6.78 lb f *ft Testing Facility The test rig that will be used for testing the fin structures is designed to test the worst case scenario by applying the force at the maximum moment. The fin test rig will be portable and consists of an L-structure that can be attached to any flat surface with a clamp. The fin structure connects directly to the test rig with the U-channel screwed into the L-structure. Static Loading Testing Various tests will be used to determine the capabilities of the test article while undergoing static loading. This is representative of the thrust during the boost phase of flight. The weight will be applied until part failure. The weight will be added in 2 lb m increments starting at 2 lb m and ending at 22 lb m. The fin must withstand the force of the weight for one minute for the run to be considered a success. Additionally, a 2.5 factor of safety shall be achieved for the testing to be considered a success imps Integrated Modular Payload System A lightweight structure is essential to maximizing payload mass fraction. Most launch vehicles use a thick walled body tube as structural member of the launch vehicle. Though simple, this design is inefficient in its use of material. The launch vehicle will be unique amongst high power launch vehicles as it utilizes an internal structure consisting of ribs and stringers as illustrated by Figure 16. These ribs and stringers are fashioned out of G of 94

22 fiberglass. The skin of the launch vehicle will consist of a minimum thickness, flexible material. Possible skin materials include fiberglass sheets, heavy-weight paper, or transparent plastic. Figure 16: CAD model of the imps To ensure the validity of the design, structural calculations were performed in order to ensure a factor of safety (F.S.) of at least 2.5 on the static loading for the structure. Consider a two dimensional view of the stringer in Figure 17. As a result of the hole, the loading path of the rod has changed significantly. That is, instead of the load being transferred through the entire crosssection of the rod, only a smaller cross-sectional area is carrying the load. The smallest crosssectional area occurs along the diameter of the fastener hole. Thus, the point of failure will be about this hole due to increased stresses. Figure 17: Dimensions for the stringer Figure 18: Cross-sectional view of the rod with the fastener hole 22 of 94

23 The smallest cross-sectional area that will carry the load is in 2 (see Figure 18). The stress concentration factor (SCF) for this particular geometry is 2, and was calculated utilizing Equation ( 3 ) SCF = A rod A min ( 3 ) Essentially, this implies that for a given load, the area around the hole will experience at least twice the amount of stress than compared to any other part. Therefore, for maximum stress of 38,000 psi 0, the maximum force that one rod can handle before breaking is lb f and was calculated utilizing Equation ( 4 ). F max = s max SCF ( 4 ) Since four stringers will be used, and the max thrust from the motor is known, the factor of safety was calculated to be 9.75, and the results are summarized in Table 13. The size of the stringers and ribs were mostly determined for adequate fastener edge clearance. This is the cause of the high factor of safety. Table 13: Values for factor of safety calculation Condition Values Max thrust from motor 408 lb f A rod in 2 A min in 2 SCF 2.05 F.S Structure Fabrication and Manufacturing Because fiberglass is a laminate composite, machining can be difficult. To simplify manufacturing, the ribs were limited to being 2-D designs. This allows the ribs to be cut from a sheet using a water-jet. Starter holes are drilled in the plate in order to prevent delamination of the G-10 when the water-jet pierces the material. The cutting jet can then be traversed from these 23 of 94

24 starter holes to the part. The stringers are cut from lengths of 3/8 inch diameter G-10 rods. They are match drilled to the ribs so that the fasteners fit well into them. To attach the skin, snaps are secured using epoxy to the outer surfaces of the holes in the ribs and to the inner surface of the skin. Structure Testing & Results Testing was performed to ensure a F.S. 3.0 on the impact energy during launch. The test rig that was used for these tests was designed to complete multiple types of test, such as static and dynamic structural loading this was done to decrease test costs through the use of a multipurpose test device as shown in Figure 19. The testing device features a rail-mounted impact machine that can hold various amounts of mass and can be lifted to various heights up to five feet for various sized test articles and/or impact energies. Figure 19: Impact testing rig Each test consisted of a known mass (3.98 kilograms) being dropped from a known distance, whose energy correlated to a design impulse (I) of 9.35 N*s. This value was determined from the calculated acceleration on the launch vehicle from the OpenRocket vehicle simulation. The height for the mass was derived using conservation of energy in Equation ( 5 ), where m t is the drop mass. Table 14 features the details of the test runs. 24 of 94

25 2 1 1 I height 2 ( 5 ) g m t Table 14: Impact test runs Test Number Impactor mass (kg) Factor of Safety Impact Energy (J) Impactor Height (m) Impactor Height (in) The test article, which consisted of half of the imps structure illustrated in Figure 20, was inspected at various locations after each run, and passed the performance criteria, with only minor damage. The results are listed in Table 15. Table 15: Testing results matrix, where X signifies damage, P signifies pass. Fastener location F.S. = 1.0 F.S. = 1.5 F.S. = 2.0 F.S. = 2.5 F.S. = p P p p P 2 P P P P P 3 P P P P P 4 P P P P P 1A P P P P X 2A P P P X X 3A P P P X X 4A P P P P P 5 P P P P P 6 P P P P P 7 P P P P P 8 P p P P P 25 of 94

26 Figure 20: imps structure test article The damage was very minor and featured discoloration from compression of the fiberglass at the faster locations in the stringers, but no fracturing occurred as seen in Figure 21. Figure 21: Evidence of minor compression damage occurring at F.S = 2.5 An internal review for this novel structural design was held to decide to continue with this option. Future testing will include full scale impact testing using a payload mass simulator Vespula Mass Breakdown Mass breakdown for each subsystems are summarized in Table 16 and Table 17 with systems level summary shown in Table 18. The values obtained for the booster section were mostly estimated utilizing Solidworks. However, since material properties were entered manually, the estimated weight should be fairly accurate. The values for nose cone, drogue chute, main chute, shock cords, imps structures, and the motor case are actual weights obtained from a scale. Note 26 of 94

27 that the ribs make up the most mass for the imps structure. This component is overbuilt structurally in order to have a satisfactory fastener edge clearance. Additionally, the mass breakdown is also presented in terms of mass fractions, as illustrated in Figure 22. Table 16: imps component weight Payload Section Weight (lbs) Quantity Total Weight (lbs) G10 Fiberglass Ribs Fiberglass Stringers " bolts Snaps Skin Total 3.27 Table 17: Booster Section Weight Budget Booster Section Weight (lbs) Quantity Total Weight (lbs) Mounting Rod Fin Tab Fin Centering Ring Empty Motor Casing Skin Fin Rear Plate Rear Cap Thrust Plate Booster fiberglass tube (6") TC Centering Ring Nuts Total of 94

28 Table 18: Subsection Mass Break Component Nose Cone 1.61 Drogue Chute + Shock Cords 1.50 Main Chute + Shock Cords 2.30 Avionics System 5.00 Payload 10.0 Payload & Recovery Structure 5.41 Booster Structure 6.62 AeroTech L Weight (lbs) Total 41.0 Figure 22: Mass Fraction for Vespula 3.9. Vespula Overall Dimension Finally, a complete CAD model of Vespula and dimensions are shown in Figure 23. Figure 23: Dimensions for Vespula 28 of 94

29 4. Launch Vehicle Performance Predictions Current mission performance predictions are based on a worst case scenario assuming a launch vehicle mass of approximately 41 pounds and an AeroTech L1390G launch vehicle motor. As the payload is finalized, mission performance predictions will be updated to reflect a more accurate mass and appropriate motor selection. For the launch vehicle to fly along the predicted trajectory, the launch vehicle must leave the launch rail at a certain velocity. For the preliminary launch vehicle design, the launch vehicle becomes stable, with a stability margin caliber of 1, at 47 feet per second, which occurs at 60 inches up the launch rail. Thus so long as our launch rail is at least 97 inches, the launch vehicle will be stable when the rail buttons leave the guide rail. 97 inches is 60 inches to reach stability plus the necessary distance between rail buttons 37 inches. The current proposed launch pad is the Apogee Gun Turret Pad. The system consists of a rail and a base. Altogether the system costs roughly $ The rail is a T-slot aluminum extrusion of approximately eight (8) feet in length and satisfies the requirements of the distance required for the launch vehicle to reach an acceptable static stability margin. Depending on how much of the budget will be left after testing and building of the initial prototype of Vespula, we may build our own launch pad to reduce cost. During recovery, the drogue and main chutes will perform as previously discussed. Table 19 takes a range of possible launch weights, without motor, for the launch vehicle and an optimal motor selection for each weight. Table 19: Best motor per launch vehicle weight and altitude reached Total Weight without Motor (lbs) Total Weight with Motor (lbs) Motor Required AeroTech L1150R-P AeroTech L850W-P AeroTech L1520T-PS AeroTech L1390G-P AeroTech L1390G-P 5259 Apogee (ft) 29 of 94

30 All simulations utilized the highest gross launch weight on the chart and the corresponding motor selection. The assumptions for all simulations are listed in Table 20, Table 20: Flight Simulation Conditions Condition Windspeed Temperature Latitude Pressure Gross launch weight Motor Value 0 mph F 34 0 N 1013 mbar 41.5 lb Aerotech 1390G-P 4.1. Flight Simulation Figure 24 shows the flight profile of the launch vehicle utilizing flight simulation conditions from Table 20. Velocity, altitude, and acceleration were plotted as a function of time. Apogee occurs at approximately 19 seconds. At apogee, the ejection charge for the drogue chute will fire, slowing the decent rate to 120 fps. Deployment of the main chute will occur at 650 feet above ground level to further decelerate the launch vehicle to 25 fps. The entire flight duration is approximately 72 seconds. Figure 24: Flight profile with AeroTech L1390 motor for a total takeoff weight of 41.5 pounds 30 of 94

31 4.2. AeroTech L1390G-P Simulated Thrust Curve The simulated thrust curve for Aerotech L1390G-P is shown in Figure 25. It is the optimum projected motor for the preliminary launch vehicle to reach an altitude of 5,280 feet. The motor will follow this thrust curve closely, however it is important to keep in mind that the performance of the motor may vary slightly in actual flight. Figure 25: Thrust curve for Aerotech L1390 motor 4.3. Stability Margin In addition, a stability analysis was performed to ensure a safe flight profile as shown in Figure 26. The stability margin of our launch vehicle during most of the flight is four calibers, where one caliber is the maximum body diameter of the launch vehicle. This is higher than the general rule of thumb among model rocketeers that the C P should be one (1) to two (2) calibers aft of the C G. However, being over-stable is not bad; it simply means that the launch vehicle will have a greater tendency to weathercock if there is any wind at launch. To counter this, our launch rod will be at least 97 inches long to ensure stability when the rail buttons leave the guide rail as mentioned previously; additional length will be added to prevent weathercock. 31 of 94

32 Figure 26: Stability margin calibers vs. Time 4.4. Drift Profile Simulation The effect of windspeeds on launch vehicle apogee is clearly demonstrated in Figure 27. The green curve illustrates our simulation data, the red triangles is to better assist the reader to visualize apogees reached at windspeed of 0, 5, 10, 15 and 20 MPH. As illustrated, there is a monotonic relationship between windspeed and apogee; the higher the windspeed, the lower the apogee. This is fundamentally due to a horizontal force created by the wind, which ultimately alters the angle of attack of the launch vehicle, and thus a lower apogee. However, from past experience with USLI, it has been shown that OpenRocket vehicle has an intrinsic tendency of over predicting the apogee. For this reason, a 3% correction factor was applied to better estimate the real result. The black line denotes our desired altitude of 5,280 feet. 32 of 94

33 Apogee (ft) MILE HIGH YELLOW JACKETS: Simulation 3% Correction 5280 ft Windspeed (MPH) Figure 27: Effect of Windspeed on Apogee To further prove this point, actual flight experiments were conducted by Mr. Niskanen, author of OpenRocket vehicle software, to demonstrate the validity of his program 0. In his experiment, two different launch vehicles were built, one utilizing B4-4 motor and the other C6-3 motor. For maximum accuracy, each component was individually weighed and the weight of the corresponding component was overridden in the software. Finally, the mass and C G position of the entire launch vehicle was overridden with measured values. Both launch vehicles were flown and the results are shown as follows, with the left figure for B4-4 motor and the right figure for C6-3 motor. Figure 28: A comparison of OpenRocket vehicle Simulation vs. actual data acquired. 33 of 94

34 The solid lines represent simulated data from OpenRocket vehicle. The upper dash lines represent simulated data from RockSim, which is truncated at apogee due to limitations of the demonstration version of the software. The lower dash lines represent the actual measured data. Clearly, both simulation software products produce optimistic data as mentioned previously, with errors in apogee about 7-10%. Thus the 3% correction is appropriate Kinetic Energy Upon Landing The kinetic energy at landing for each independent and tethered section of Vespula was calculated utilizing Equation ( 6 ) where m is the mass of each section and v is the velocity. The results are summarized in Table 21. KE = 1 2 mv2 ( 6 ) Table 21: Kinetic energy upon landing for each section of Vespula Mass (lbs) Velocity (ft/sec) KE (ft-lb) KE (joules) Nose Cone Booster Section Payload Section of 94

35 5. Launch Vehicle Testing 5.1. Subscale Testing A subscale flight test will be performed prior to CDR to determine the feasibility of certain aspects of our design. This testing primarily focuses on the performance of our exterior skin covering the booster and payload section during flight. The basic design of our subscale launch vehicle, to be called Korsakov features a smaller diameter body tube, that is covered by a nonload bearing, thin-walled skin. Korsakov will utilize two (2) 0.5 inch diameter launch lugs, one placed on the top section and one on the bottom. They will be two (2) inches long and epoxied onto the sides. The thrust is transfer through the internal cardboard tube from the booster section to the nose cone. Due to the configuration of the launch vehicle, a dummy mass of 0.3 pounds was added right after the nose cone to provide at least one (1) caliber of stability margin (see Figure 29). The characteristic of Korsakov is summarized in Table 22. Figure 29: Korsakov launch vehicle layout Table 22: Characteristics of Korsakov vehicle Condition Length Outer Diameter Value 45 inches 3 inches Motor Aerotech H242T Max Mach Number 0.44 Again, a stability analysis was performed to simulate the C G and C P translations in flight as shown in Figure 30. The stability margin of Korsakov is around 2.4 calibers for most of the flight, thus it will be stable. 35 of 94

36 Figure 30: Stability profile for Korsakov vehicle Table 23 shows all the materials required for building Korsakov with a total projected cost of $ The total weight was estimated to be 3.4 lb f utilizing OpenRocket vehicle software. 36 of 94

37 Table 23: Material and cost for Korsakov Item Use Cost ($) Motor reload kit Launch vehicle motor Airframe Tube 74/18 (Thin Wall 3" tube) Main body Airframe Tube 29/13 Inner body Airframe Tube 56/18 (Estes BT-70 size) Motor tube Baltic Birch Plywood 6mm-1/4" x 24" x 30" Fins /4" plywood Centering rings U-bolts Recovery Poster board Skin Snaps Skin fasteners Parachute Recovery Nose cone Aerodynamics Total of 94

38 6. Launch System and Platform As mentioned previously, the launch system that is showing the most promise is the Apogee Gun Turret Pad. The system consists of the rail and the base. Altogether the system costs $ The rail is a T-Slot aluminum extrusion of approximately 8 feet in length and satisfies the requirements of the distance required for our launch vehicle to reach an acceptable static stability margin. The blast deflection pad is angled with dimensions 9 inches x 9 inches x 0.25 inches, and is made of heavy-gauge steel. The rail and the deflection pad are attached to the head of the base, which can pivot horizontally for easy loading of launch vehicles. Three legs with leveling screws are attached to the base so that the launch angle can be adjusted to desired conditions, with all parts also being constructed from heavy-gauge steel. The entire system weighs roughly 30 pounds, and is collapsible for transportation. Our launch vehicle will have a two rail buttons attached such that they do not interfere structurally with any other components. The rail buttons slide into T-shaped aluminum extrusion and limit the launch vehicle s motion except in the desired launch direction. The first button will be attached to the thrust plate in the booster section, where components exist to easily create an attachment interface with minimal structural interference. The second button will be attached the rear thrust retention plate. This will ensure that both buttons will be on the rail for a time, and at least one button will be on the rail for the length required of the launch vehicle to reach an acceptable static stability margin upon launch. The launch procedure checklist is presented in Appendix I. 38 of 94

39 7. Payload 7.1. Introduction to the Experiment and Payload Concept Features & Definition Today, many entrepreneurs are beginning to build newer and more cost-effective launch vehicles. Every one of these launch vehicles must address a specific challenge in their design process: integration with the spacecraft payload. This integration presents difficulties in launch vehicle design because harmonic oscillations of the spacecraft mass could cause structural damage to either the launch vehicle or the spacecraft itself. To solve this dilemma, industry typically utilizes large mechanical springs in addition to the placement of certain structural constraints on the payload spacecraft for use of a particular launch vehicle. Repeated deformation on vibration dampers and springs used in launch vehicles presents a further issue in providing reusability, as these parts must be intermittently replaced. Furthermore, modifications must be made to both payload and launch vehicle to tune the natural frequencies of both and prevent harmful oscillation. The net result of the present situation is an increase in overall structural mass, which combined with the necessary increase in fuel required and maintenance, dramatically increases the launch cost to the detriment of mission capability. The Mile High Yellow Jackets intend to provide a possible alternative solution in a demonstration of the ability of electromagnetic levitation to lower the necessary structural masses currently required to prevent harmonic oscillation, decreasing launch cost Overview of the Experiment Hypothesis and Premise The premise of the experiment is that If a platform can be levitated and stabilized in a dynamic magnetic field during the flight of a launch vehicle, then greater stability and lower structural mass may be achieved. The payload will utilize dynamic three-dimensional magnetic fields to create an Active Platform Electromagnetic Stabilization, or A.P.E.S., system for use during the ascent phase of the launch vehicle s trajectory. The launch vehicle ascent will provide a high vibrational intensity 39 of 94

40 environment in which to test the stabilization system. Two further premises are necessary for this A.P.E.S. system, namely: 1. Under appropriate conditions, it is possible to control complex oscillating magnetic fields such that a system of ferromagnetic materials or permanent magnets may be levitated in non-rotational stability. 2. A design exists such that a platform of some size and low mass may be levitated using ferromagnetic materials or permanent magnets. Therefore, after completing a thorough analysis of the dynamics of materials being levitated and stabilized in magnetic fields, the will implement a design to apply this science to a platform within the Vespula launch vehicle. Experimental Method and Relevance of Data The parameters to be measured in the experiment are the coordinates of the position of a test sample in the case of ground testing and the coordinates of the platform for the ascent of the launch vehicle. To ensure full implementation of the scientific method, the experiments will be carried out such that the results are controlled by comparison i.e. a series of tests of increasing complexity will be conducted such that the control theory of the A.P.E.S. system is constructed methodically. Analysis of optical and infrared data will provide Cartesian coordinates, which, in a known design, can be used to specify the position and displacements of the platform, defining all kinematics to a level of accuracy proportional to the sample and computational rates of the data acquisition system. This data provides an empirical basis for confirmation or rejection of an experimental hypothesis a newly written control program can be considered as a hypothesis and for the improvement of the system as a whole. 40 of 94

41 Figure 31: Section view of the proposed flight model of the A.P.E.S. system Ground Test Plan Ground testing will serve two general purposes: (1) the development of data algorithms and control laws, and (2) the verification and validation of theory and control systems. By understanding and predicting the kinematics of a sample plate driven by dynamic magnetic fields, steady-state models can be formed allowing for stable non-rotating levitation with oscillation dampening. The testing process will ramp the complexity of the model, beginning with simple 1-dimensional tests, and increasing the number of solenoids and dimensions until a 3-dimensional multi-solenoid model has been created. Data collected from ground testing will directly inform the control of the flight A.P.E.S. system design. The flight design is depicted in Figure 31, and the ground test platform design is illustrated in Figure of 94

42 . Figure 32: The A.P.E.S. system ground test platform Test Sequences Ground testing will undergo test sequences. Preliminary testing will calibrate instruments for use in the experiments. The first sequence will simply lift a ferrite toroid ring within a tube, constricting motion to a single axis. The ring will be lifted to an equilibrium position. This test will provide a validation of the force equations moving forward. Subsequent tests will increase the number of dimensions and the number of solenoids involved. In each case, an equilibrium position will be sought. For two- and three-dimensional testing, computer vision will be used to provide kinematics data for controlling solenoid current. An Arduino Mega may be used for simplicity during ground testing; however, some team members do possess an ARM Cortex M3 Blueboard, so this is also an option for control of solenoids and receipt/analysis of sensor feedback. The preliminary testing plan is included in Appendix 2. Mathematical and Physical Modeling of Magnetic Fields In order to accomplish the objective of stabilizing a platform with magnetic fields during the ascent of a launch vehicle, a control system must be developed with inputs of voltages and currents supplied to solenoids and optical sensing feedback for kinematics data. To create the control system, equations and experimentation to model the fields and resultant forces on an 42 of 94

43 object in the field will be derived and conducted, respectively, from the scientific principles governing electromagnetism. Typically, electromagnetic equations are focused on defining axial interactions, while the A.P.E.S. experiment requires a comprehensive understanding of threedimensional magnetic fields. The following sections will define the governing equations and concepts that are the foundation for the experimental testing and will serve as the basis for a data-centered control system. Modeling General Magnetic Fields If two magnets or electromagnets are at a large enough distance from each other, or small enough compared to the distances involved, then they can be modeled as being magnetic dipoles. A magnetic dipole can be thought of as a small current loop; this still creates a non-vanishing magnetic field at distances much larger than the radius of the loop. The magnetic dipole moment of a single current loop is defined as (1) where the S vector, and hence m as well, is oriented perpendicular to the planar area of the loop so that curling the fingers of one s right hand in the direction of the current gives the direction of S as the direction of the thumb. The magnetic potential due to a magnetic dipole of moment m is ( ) (2) where r is the vector from the dipole to the field point where the potential is being calculated, r is the magnitude of vector r, and μ is the permeability of the medium at the field point. The magnetic flux density B and the magnetic field H due to the dipole are, respectively, ( ) ( ( ) ) (3) ( ) ( ( ) ) (4) Where is the unit vector in the direction of r, and the distance r is much greater than the radius of the loop. 43 of 94

44 There are two ways to approximate model the vector potential field, the magnetic field, and the magnetic induction field as produced by a solenoid using these equations. The first method is to model the solenoid as a single dipole of moment at the center of the solenoid, where N is the number of turns in the solenoid, as a solenoid has N current loops each of moment IS. However, this does not take into account the fact that each loop of the solenoid is not at the same location. Therefore, a more precise way of modeling the solenoid albeit still an approximation would be to place one dipole of moment IS at the center of each loop that makes up the solenoid, or perhaps one moment per k loops of moment kis, where we have a choice of k. However, computational difficulty is greatly increased due to the necessity of finite-element solver techniques as the mathematics progresses. The magnetic H field produced by each model are shown below, where N is taken to be 11 loops (distributed over 2 cm of length for the second model) and is taken to be k. One dipole of moment 11k is placed at the origin in the typical cartesian plane in Figure 33, and 11 dipoles of moment k distributed from -1 to 1 along the y-axis in Figure 34, for the sake of simplicity. are Figure 33: field generated by a single dipole 44 of 94

45 Figure 34: field generated by multiple dipoles Generation of Magnetic Forces in Materials All materials are composed of atoms, with a positively charged nucleus and negatively charged electrons. The movement and rotation of these charges form microscopic magnetic dipoles, which have magnetic dipole moments. The magnetization vector, M, of a material at a point is defined as the volume density of magnetic dipole moment, i.e. (5) where each is the magnetic moment of the kth atom in the volume, and the sum is over all the atoms. The force on a magnetic material can be determined by summing the forces on the dipoles in the material due to the field that it is placed in. The magnetization of a material depends on the field it is placed in, and the flux density depends on the field, as follows: (6) ( ) ( ) (7) where is the material s magnetic susceptibility, is its relative permeability, and is the absolute permeability. The parameters and are not always constant, especially in the case of ferromagnetic materials. However, assuming a linear relationship between M and H approximately true in the case of magnetically soft ferrite or a constant M in the case of a 45 of 94

46 permanent neodymium magnet, using the H field of a dipole or multiple dipoles as the field of the solenoids, the force on the platform due to the fields interacting with the microscopic dipoles in the material can be calculated. Forces on Materials in Magnetic Fields The force on an object is the sum of the forces on all of the magnetic dipoles that make up the object. By definition, the magnetic dipole moment of an infinitesimal volume of the object dv is. The force due to the field of a magnetic dipole of moment on a magnetic dipole of moment m that is in a material of permeability is: ( ) [( ) ( ) ( ) ( )( ) ] (8) Where r is the vector from to m, and is again the unit vector in the direction of r. First the case of a ferrite platform is considered, with approximate constant and μ. In this case, the force on the platform is calculated to be: ( ) [( ) ( ) ( ) ] (9) Where is now used as for the solenoid and the integral is evaluated over the volume of the platform. If it is assumed that the object is small such that the quantity integrated does not vary significantly over the volume, the force on the platform of volume V, due to the solenoid of moment, is: ( ) [( ) ( ) ] (10) Where is the unit vector in the direction of S the unit normal to the loop area of the solenoid and r is the position vector from the solenoid center to the center of mass of the platform. While approximate, it is clear that the force will vary as the square of current and inversely by the seventh power of the distance between the solenoid and the platform assuming a magnetically-soft ferrite material. To check the validity of this equation, and assuming that both 46 of 94

47 and r are in the positive k direction in a Cartesian plane, such that the platform is above the dipole, it is found that: (11) Or that the platform is pulled towards the dipole, which matches the basic experience of magnetic materials attracted to magnets due to induction. The equations given above are derived in Appendix 3. However, the validity of these equations is primarily for the case of a single solenoid acting on a platform with constant permeability. Forces originating from more than one solenoid do not add in the conventional sense, as the induction of a ferrite material is highly non-linear. These equations must be re-derived from equation (8), as the fields and magnetization of the platform change in the n-solenoid problem. Much easier is the case of a permanent neodymium magnet with constant magnetization M throughout. In this case, the force on the platform is the sum o0f the force on each segment, ( ) [( ) ( ) ( ) ( )( ) ] (12) Here, the constant involves rather than just, since the M vector is constant and is largely independent of H. Again, the exact value of the expression is highly dependent on the shape of the volume integrated upon. However, if the volume V is small, the force can be taken due to the solenoid field NIS as: ( ) [( ) ( ) ( ) ( )( ) ] (13) Where is defined as before. Equation (11) is also an approximate solution, but here it is evident that the force on a permanent magnet varies only directly on the current in the solenoid and inversely by the fourth power of the distance, rather than by the square of current and 47 of 94

48 inversely by the seventh power of distance in the case of forces from induction in a ferrite platform. The force will also depend on the orientation of M. Unlike for the case of a material with constant permeability, the forces on a permanent magnet due to multiple solenoids do add in the conventional sense, greatly simplifying computational analysis. Scientific Merit and Success Criteria Scientific Merit The problem of magnetic force interactions from n-solenoids on a single sample is a non-trivial problem in electromagnetics. The difficulty in describing complex field relationships is similar to the difficulty in aerodynamics for describing complex fluid flows, and many of the computational techniques are similar. However, due to the nature of the complexity, the study of complex magnetic interactions must be a data-driven process, as in aerodynamics. The A.P.E.S. system will depend upon a theory-informed, data-driven model for control. This data will be generated through a series of ground test experiments that gradual increase the complexity of the problem. Final model testing on the ground will involve only permanent magnets and solenoids, simplifying the force interactions to compensate for complex geometry. The A.P.E.S. project may be considered as a dual scientific-engineering payload. A period of scientific analysis is necessary, as stated above. However, the actual product flown in the launch vehicle will be flown for verification and validation purposes after the conclusion of ground testing; the flight will test the performance of the derived model, and engineering design, during the dynamics of the ascent phase. This process of scientific investigation followed by engineering development is not entirely unlike the development of experimental aircraft and spacecraft, where some scientific investigation may be needed before the engineering can proceed Experimental Requirements and Objectives Included within the terminology of Flight Systems is included all avionics and experimental material necessary for a successful mission. Therefore, the Flight Systems group must deliver based on two sets of functional requirements. Requirements for 48 of 94

49 A.P.E.S. are detailed in Table 24. These four basic requirements have been updated for functionality from those originally listed in the proposal and all elements necessary to the smooth operation of the experiment are tied to these requirements. Table 24: A.P.E.S. system requirements Requirement Number Requirement Definition 1 A platform shall be levitating without rotation during the ascent of the launch vehicle and coordinate data provided through optical and infrared sensing shall be used as verification metrics. 2 All elements of the A.P.E.S. system shall weigh no more than 10 pounds total, including all hardware and electronics in direct support of the A.P.E.S. system. 3 The A.P.E.S. system must actively correct for accelerations of the launch vehicle. 4 The A.P.E.S. system must shut down safely at apogee. There are several options available to Flight Systems to satisfy the functional requirements. The use of the word platform is necessarily open-ended as the team investigates several types of platforms. Ground testing requirements necessitate that several platform configurations be used to formulate the final system model. Table 25 outlines some of the ground testing objectives for A.P.E.S. Table 25: A.P.E.S. ground testing objectives Objective Number Objective Definition 1 Ground testing should develop a data-driven model for the interaction of ferromagnetic materials and permanent magnets in one-, two-, and threedimensional dynamical systems of magnetic fields. 2 Ground testing should test the feasibility of de-scope and full flight designs. 3 Ground testing should provide the computational basis for control of the A.P.E.S. system during ascent of the launch vehicle. If ground testing results are successful and combined with a thorough theoretical understanding of the phenomena involved in dynamic magnetic fields, ground testing will enhance the ability of Flight Systems to successfully accomplish the most important objective of the A.P.E.S. system that the A.P.E.S. platform shall not touch any part of the container during the ascent phase. Optical sensing of either the infrared or visible spectrum will serve to inform the control system of the location of the platform. A full flight model illustrated in a section view in Figure of 94

50 would include a series of concentric cylinders, wherein the center cylinder would contain permanent magnets and a platform, while the inner cylinder would store solenoids and a CMOS camera for optical detection of fiducial markers. Success Criteria A fully successful launch of the A.P.E.S. system will demonstrate a capability to respond and dampen impulses delivered to the levitated platform from the launch vehicle during ascent. Namely, the platform shall, ideally, not touch any other A.P.E.S. components during ascent. However, failure to achieve total isolation of the platform is only a partial failure, as conditions during launch could lead to unforeseen perturbations of the launch vehicle. Any failure to levitate the platform, dampen motion, shutdown, or start up, would be a critical failure. Specified preliminary success criteria are given in Table 26: Success Criteria. An attempt has been made to provide a reasonable goal for system performance before any physical testing has begun. It is not certain whether it will be possible to provide the same level of control for many disturbances. Table 26: Success Criteria Success Criteria The A.P.E.S. system shall be successful if 1 For an impulse response the system shall have a 2% settling time of less than 0.5 seconds 2 And for the flight of the launch vehicle that the platform should not touch other components of the A.P.E.S. system. 50 of 94

51 8. Flight Systems 8.1. Flight Avionics Overview and Requirements To successfully complete the USLI mission, flight systems is further responsible for providing a fully functional flight computer system. Avionics is the second subsystem of Flight Systems, responsible for data acquisition, experimental control, recovery electronics, and features necessary as per the USLI Handbook. Avionics requirements are listed in Table 27. Requirements and Products Two major products of the Avionics subsystem are the flight computer and the experiment computer for A.P.E.S as detailed in Figure 35. The flight computer interfaces with all sensors not directly involved with A.P.E.S. and controls most sensing, logging, and telemetry for the launch vehicle. The A.P.E.S. computer focuses entirely on the control of the A.P.E.S. system. Figure 35: General products of Flight Avionics 51 of 94

52 The A.P.E.S. computer will calculate the position of the platform and control the solenoids in order to change the magnetic field and stabilize the platform. Independent computing systems provides modularity for ease of implementation and debugging. The methodology for component selection shall include consideration of clock speed, I/O, and voltage requirements. Electromagnetic interference will be shielded by a Faraday cage. The system will incorporate redundancy to tolerate the loss of one or more sensors and/or communication lines. Table 27: Flight Avionics Requirements Requirement Number Requirement Definition 1 The avionics shall be a thoroughly tested custom design. 2 The flight computer shall be based off of the ATMEGA 2560 chip. 3 The A.P.E.S. computer shall be a commercially purchased board based upon the ARM Cortex M3 microprocessor. 4 The flight avionics shall collect data on acceleration, altitude, and data necessary to the A.P.E.S. payload and execute the A.P.E.S. control laws. 5 Key elements of the flight systems shall operate on independent power supplies. 6 Power supplies should allow for successful payload operation during launch vehicle flight with up to 1 hour of wait on the launch pad and 2 hours of wait during launch vehicle preparation. 7 The flight avionics, with the exception of the recovery avionics, shall begin at launch and these systems should be capable of being armed externally to the launch vehicle structure. 8 The avionics should provide for the communication of ground location to base station for recovery. 9 Separate elements shall trigger ejection charges for parachutes, and a competition altimeter shall be included. 10 GPS coordinates of all independent launch vehicle sections shall be transmitted to a base station. 11 The recovery avionics and system shall be separate from the main flight avionics. 12 The base stations shall be capable of receiving and displaying data transmitted from on board the launch vehicle. Flight Computer The flight computer will run the ATMEGA 2560AU processor with the Arduino bootloader and other necessary components for ease of programming. The chip has sufficient I2C, serial, and analog inputs to read data from all sensors and log to an SD card based on Sparkfun s OpenLog break-out board. Additionally, the chip will run the Fastrax UP501 GPS module and send the data to an Xbee PRO for transmission to the ground station. An OpenLog board will provide 52 of 94

53 logging capabilities. The chip will be programmed in the Arduino language, a subset of C++ with some additional libraries. Figure 36 provides a generalization of proposed flight computer software. Figure 36: Generalization of flight computer software The flight computer must accomplish several tasks and handle multiple responsibilities. The main goal of this system is to collect and monitor all the relevant data from the environment around it such as the strain on the launch vehicle, environmental factors such as temperature, stray magnetic flux from the A.P.E.S system, launch vehicle acceleration, and GPS position. During flight, the flight computer must also monitor the payload's control system and data through a serial bus and provide an emergency secondary disengage for the A.P.E.S. system in the case of a necessary emergency shutdown. During flight the avionics will log all data to a SD card. Solid state memory should allow recovery of flight data if a recoverable failure occurs. Post-recovery, the flight computer must switch to location and communication systems to transmit a GPS signal through the telemetry system to the ground station. Figure 37 and Table 28 provide the current proposed flight computer schematic and major components, respectively. 53 of 94

54 Figure 37: Proposed layout of flight computer. Larger copy included in Appendix 4 54 of 94

55 Table 28: Major Flight Computer Components Part Number Component Picture Description 1 The flight computer microprocessor, the ATmega The GPS receiver, the Fastrax UP501 GPS module 3 The Xbee PRO 900-XSC module for communication between launch vehicle and ground station 4 The OpenLog board will provide logging capability A.P.E.S. System Computer The A.P.E.S. system and its computer will focus on the stabilization of the isolated platform. The computer system will be a commercially purchased ARM Cortex M3 breakout board named a Blueboard. The Cortex M3 will provide up to 100 MHz of clock speed for the rapid computation necessary in analyzing optical data. The Blueboard will also allow the Cortex M3 to handle all control of the magnetic fields through pulse-width modulation (PWM), and delivery of data to the flight computer. A.P.E.S. system software will also include steps to shut down the 55 of 94

56 experiment at or around apogee, and physical constraints will be built into the A.P.E.S. system structure to improve shutdown and recovery safety. An outline of the A.P.E.S. system software is given in. and a rough layout of A.P.E.S. computer hardware is given in Figure 39. Figure 38: Generalization of A.P.E.S. system software Figure 39: Layout of A.P.E.S. computer system hardware 56 of 94

57 8.2. Power Systems Power Supply The avionics system, including computers and sensors, will be powered by a 9V battery. The supply will be attached to the Avionics Computer Board which is designed to have a voltage regulator circuit providing 3.3V and 5V rails. The supply will provide 1200mAh of power. Figure 40: Discharge characteristics of the A123 battery The avionics use a negligible amount of power in sleep mode providing a minimum of 5 hours of wait capability for the launch pad. Upon launch the system is activated and will have greater then needed power capacity to perform its duties until the launch vehicle is retrieved. Separately, the A.P.E.S. system will utilize a four-pack of A123 lithium iron phosphate (LiFePO) rechargeable batteries, one of which is shown in Figure 41. These batteries have a per-unit nominal capacity and voltage of 2.3 ampere-hours and 3.3V, respectively. Furthermore, the A123 batteries provide a maximum discharge rate of 70 amperes. Figure 40 illustrates the discharge characteristics of the A123 at four discharge rates. The ability of the A123 to provide a large current is critical to the A.P.E.S. system, which will rely on pulse-width modulation to change magnetic field intensity via manipulation of a root-mean-square current. 57 of 94

58 Figure 41: A single A123 LiFePO battery 8.3. Telemetry and Recovery Ground station The ground station for receipt of data shall consist of a laptop connected via USB to an Xbee Pro and Xbee Explorer with a rubber duck antenna. This will ensure simplicity, portability, and operability of the ground station. Transmitter Design In order to satisfy recovery requirements that the launch vehicle be found, a GPS module is included with the avionics in addition to radio communication equipment. The Fastrax UP501 provides a 10Hz update rate, rapid satellite acquisition, and low current draw. Position data is logged on-board and transmitted over the 900MHz radio band to our ground station. The telemetry system is designed to utilize two Xbee PRO 900-XSC modules for one-way communication from the launch vehicle to the ground station. Using a simple, loss-tolerant protocol with reliable delivery ensures the data is received if at all possible and that the information is correct. To extend the range beyond 1 mile, each module has a 900MHz monopole-monopole vertically polarized rubber duck antenna with 2 dbi gain and 10W of power. This antenna s performance is depicted graphically in ory.. Receipt of GPS data via radio to the ground station will satisfy the recovery requirement and bolster kinematics data of the launch vehicle trajectory. 58 of 94

59 Figure 42: Antenna performance as a function of range 8.4. Integration Modularity and Motivation The modular internal launch vehicle structure permits integrating the payload with minimal effort. A section of the internal launch vehicle structure is reserved for the altimeters, payload experiment, and flight computer. The ends of the modular section are fitted structurally with solid fiberglass to sustain the bursts of the recovery system. These solid fiberglass sections are fitted with shear pins to maintain stability during flight. The entire system stacks together for dual deployment. The volume designed for the payload is an area between the struts 8 inches long with an average diameter of 3.2 inches within the hexagonal inner side of the ribs. The A.P.E.S. device will be anchored to the top of this section via a "universal bracket" to one of the ribs, shown in. Below A.P.E.S. will be the flight system computer will in a shielded compartment. The computer will also be mounted to a rib using a universal bracket. 59 of 94

60 Figure 43: Universal mounting bracket bolted to rib Universal Mounting Bracket The unique and robust structure of the Vespula launch vehicle will allow greater reusability and modularity for integration of the flight and payload subsystems. In order to speed up integration time a generalized universal mounting system will be employed so that current and future subsystems can be quickly and effectively mounted to the major rib sections of the launch vehicle structure. When these and future internal components are designed and produced, minimal design consideration will be needed to account for attachment to the universal mounting bracket. This continues on the Georgia Tech tradition of simplifying and unifying the structural elements of the launch vehicle, simplifying design and improving construction time and structural robustness. The universal mounting bracket shall be built to accommodate structures such as the A.P.E.S. device with minimal fabrication and design requirements, and to fit within rib structures with few necessary design parameters on the ribs themselves, increasing the potential reusability of the universal mounting bracket should future teams alter the modular structure further. As shown in Figure 31, the proposed A.P.E.S. structure mounts easily with the universal bracket. The use of the bracket also allows for the whole A.P.E.S. system to be removed more easily in case there any modifications are required. Furthermore this frees up the designs for the payload structure as it is fundamentally non-load bearing as the internal launch vehicle structure is taking care of that already. 60 of 94

61 Figure 44: Basic finite-element-analysis of the universal mounting bracket Finite-element-analysis carried out in the Solidworks Office environment details an axial load of 100 lb f placed in the center of the mounting bracket. High stress regions appear due to the fallibilities of the Solidworks finite-element-analysis software and the difficulty of specifying distributed reaction loads. The results of the finite-element-analysis appear in Figure 44. While quarter inch (0.25 ) aluminum was necessary to provide a factor of safety of 2 in the case of point application of resistive loads, it is more likely that 7/32 or 15/64 thickness of material shall suffice. Further study will be done to specific an exact safety factor and the design may be modified to ensure a minimum safety factor of 2. The universal mounting bracket will mount into the launch vehicle structural ribs at approved attachment points using number 8 bolts. Initial designs allow for the A.P.E.S. structure to then bolt directly to the bracket also using number 8 bolts. Table 29 provides further characteristics of the universal mounting bracket. Performance evaluation metrics will be developed further as elements of Flight Systems begin to be produced. However, several basic metrics exist already, namely, the reduction of all motion 61 of 94

62 of the A.P.E.S. plate in a timely manner, i.e. a well-damped impulse response. The flight computer must survive for several hours on the launch pad. All data must be handled and recorded accurately by both the flight and experimental computers. Table 29: Universal Mounting Bracket Specifications Universal Mounting Bracket Parameter Thickness Diameter Bolt Hole Diameter Rib Mount Bolt Radius from Center Number of Bolts to Rib A.P.E.S. System Mount Bolt Radius from Center Avionics System Mount Bolt Radius from Center Mounting Bracket Material 0.25 in in in in. 8 bolts maximum in in Aluminum Design Value 8.5. Sensing Capabilities Flight Avionics Sensors General Sensing Sensing data will be provided to the flight computer through ten (10) different sensors chosen to give the most relevant data regarding experimental and launch vehicle performance. The sensing system must account for difficulties arising in communication interfaces, voltage requirements, and material sourcing. Components must survive periods of potentially strong magnetic flux density and sensors placed in launch vehicle modules which undergo separation must resist explosive impulses which may interfere or damage sensing components. Kinematics and Location The accelerometer ADXL345 (shown in Figure 45) will provide acceleration data and, combined with the GPS module, provide rotation and position data for the launch vehicle trajectory. Three axis capabilities will implicitly define velocity, position, and rotational motion. The ADXL345 accelerometer can record up to ±16G. The ADXL345 is capable of entering a standby mode for periods of inactivity, an advantage for periods of inactivity during setup and preparation to launch. 62 of 94

63 Figure 45: ADXL345 accelerometer Magnetic Fields The chosen magnetometer for detection of potentially harmful fields in the vicinity of the avionics is the HMC1043, and will determine the effectiveness of our shielding to contain the magnetic fields from the APES system. The sensor detects the magnetic field in three dimensions and static testing will allow for compensation for the contributions of the Earth s magnetic field. The sensor will be used to determine the flux within the avionics bay of the launch vehicle to help monitor the influence of the magnetic field to our other equipment. The HMC1043, in 63 of 94

64 , can sense up to ±6 gauss. A combination of distance and mu-metal shielding should diminish the A.P.E.S. fields significantly that such a small range should be appropriate. 64 of 94

65 Figure 46: HMC1043 Magnetometer A.P.E.S. System Sensing The A.P.E.S. computing system will require two types of sensors for feedback and control. Position of the levitating test platform inside of A.P.E.S. will be tracked and displacement data used for derivation of platform kinematics. The magnetic fields generated by the solenoids will also be monitored and compared to models and thresholds developed during ground testing. There are several serious issues to sensing in the A.P.E.S. experiment. First is the possibility of large magnetic flux potentially as large as several hundred gauss. Strong magnetic flux will induce current in wiring often destroying sensitive digital electronics. High current such as the solenoids power cables may also risk induced current in electronics. To counteract these issues, mu-metal Faraday shielding and distancing from the computational elements will be utilized. However, sensors and detection equipment will be chosen to satisfy the expected parameters of the environment around the A.P.E.S. system. 65 of 94

66 Table 30: Possible A.P.E.S. distance sensors Sensor Cartesian Coordinate Axes Viewing Angle >45 Degrees Range <5cm Resolution <1mm Delay <20ms Interfere nce Flux Sensitive Reliable under Shock and Vibrations Small Form Factor <15mm Ultrasonic Distance 1 No No No No No No No No IR Distance 1 No Yes Yes Yes No Yes, w/shield ing Yes Yes Laser Distance 2 Yes Yes Yes No No Yes, w/shield ing Yes No CMOS Camera 2 Yes Yes Yes Yes Yes Yes, w/shield ing Yes Yes A.P.E.S. Distance Sensing The full three-dimensional design would utilize two to three CMOS cameras. The OVM7690 Camera Cube CMOS camera meets all current design requirements and expected environmental conditions, as outlined in Table 30. The camera sensor is a small form factor (2x2x1mm) color image camera module, as illustrated in Figure 47, with integrated optical glass lens and on-chip image processing. A test pattern is used for initial software pixel-to-distance mapping and calibration for the camera output data and is mounted on the opposite side of the test structure as the camera. The test pattern must be an easily identifiable pattern this will be made of fiber optic cables mounted on a panel attached to a specific color light emitting diode (LED). A second camera and test pattern are mounted perpendicular to the first, using a second specific color. 66 of 94

67 Figure 47: OVM7690 Camera Cube Once the calibration is complete, the cameras output is read at frames per second (FPS). The levitating platform is painted another specific color which is not the same as the colors already used for calibration patterns of the two cameras. Several white LEDs are used for flooding light to make the painted platform visible. A simple edge detection algorithm is used to find the displacement of the platform in each sample frame. Using the pixel to distance mapping from the initial calibration, the displacement between samples is calculated which in turn is used to update the Cartesian coordinate for the platform in three dimensions. The use of one camera for sensing on three axes was rejected because the small movements along the forward line of sight (depth perception) would not be interpreted with high enough resolution. Movement along the horizontal and vertical plane will be sufficient, so two axes can be used. A.P.E.S. Magnetic Field Sensing While not critical to the flight mission of the A.P.E.S. system, sensing of magnetic fields during ground testing of the A.P.E.S. system will allow for better model generation as well as full confirmation of the theoretical basis of the project. Therefore, the MLX90363 magnetometer has been proposed for use in ground testing applications. The sensor is sensitive up to Tesla, has a sample rate of 1 millisecond, and outputs magnetic field direction as a three coordinate vector. This sensor was chosen for its high magnetic flux sensitivity, decent resolution 67 of 94

68 increments, and three axis direction vector output. The magnetometer controls on-chip digital signal processing De-Scope Options Payload De-Scope The de-scope option for the payload will utilize a single infrared distance detection along a single axis of linear motion, where a sample plate is levitated vertically at constant distance between the sensor and a solenoid during the ascent of the launch vehicle. The objective of the flight experiment would be to then maintain a constant position with a 2% settling time of less than 0.5 seconds for a unit impulse. This option provides an opportunity to develop many of the same control theories without the necessity to grapple complex 3-dimensional magnetic fields. Flight Computer De-Scope The de-scope option for the flight computer is to use a commercially available Arduino Mega board, rather than fabricate a custom board. This option provides full capability should timelines not permit the completion of design and construction of the custom board. 68 of 94

69 9. Safety 9.1. General Safety Ensuring the safety of our members during building, testing and implementation of the payload experiment is an ideal condition. Procedures have been created and implemented in all of our build environments to ensure safety requirements are met and exceeded. A key way the Yellow Jackets ensure team safety is to always work in teams of at least two when using equipment or during construction. This guarantees that should an incident occur with a device the other member could provide immediate assistance or quickly get addition help if required. The Invention Studio where the team does a majority of its work is equipped with safety glasses, fire extinguishers, first aid kits, and expert personnel in the use of each of the machines in the area. All the members of the payload and flight systems teams have been briefed on the proper procedures and proper handling of machines in the labs Payload Hazards As already mentioned in General Safety, the same methodology to identify and assess risks for vehicle safety will be used to identify hazards for the payload. The entire payload and flight systems teams have been briefed on the possible hazards they may encounter while working with the payload and how to go about avoiding them. Hazards that relate specifically to the payload are listed in Table 31. Payload failure modes are outlined in Table 32. Table 31: Hazards, Risks, and Mitigation Hazard Risk Assessment Control & Mitigation Electrocution Serious Injury/death Do not touch wires that are hot and not insulated. Wear rubber gloves when the device is in operation. Handle leads to the power supply with care. Use low voltage settings whenever possible. Electromagnetic Fields Interfere with electronic devices inside the body Ground test equipment, keep people with electronic components in them away from the coil when the electromagnetic coil is in use. 69 of 94

70 Hazard Risk Assessment Control & Mitigation Epoxy/glue Toxic fumes, skin irritation, eye irritation Work in well ventilated areas to prevent a buildup of fumes. Gloves face masks, and safety glasses will be worn at all times to prevent irritation. Fire Soldering Iron Burns, serious injury and death Burns, solder splashing into eyes Keep a fire extinguisher in the lab. If an object becomes too hot or starts to burn, cut power and be prepared to use a fire extinguisher. Wear safety glasses to prevent damage to eyes. Do not handle the soldering lead directly only touch handle. Do not directly hold an object being soldered. Drills Dremel Serious injury, cuts, punctures, and scrapes Serious injury, cuts, and scrapes Only operate tools under supervision of team mates. Only use tools in the appropriate manner. Wear safety glasses to prevent debris from entering the eyes Only operate tools under supervision of team mates. Only use tools in the appropriate manner. Wear safety glasses to prevent debris from entering the eyes Hand Saws Cuts, serious injury Only use saws under supervision of team mates. Only use tools in the appropriate manner. Wear safety glasses to prevent debris from entering the eyes. Do not cut in the direction of yourself or others. Exacto Knives Hammers Power Supply Cuts, serious injury, death Bruises, broken bones, and serious injury Electrocution, serious injury and death Only use knives under supervision of team mates. Only use tools in the appropriate manner. Do not cut in the direction of yourself or others. Be careful to avoid hitting your hand while using a hammer. Only operate power supply under supervision of team mates. Turn of power supply when interacting with circuitry. 70 of 94

71 Hazard Risk Assessment Control & Mitigation Batteries Explode Eye irritation, skin irritation, burns Wear safety glasses and gloves. Make sure there are no shorts in the circuit. If a battery gets too hot stop using it an remove any connections to it. Improper Dress during construction Exposed construction metal Neodymium Magnets Serious injury, broken bones Punctures, scrapes, cuts, or serious injury Pinching, bruising, and snapping through fingers. Wear closed toe shoes, clothing that is not baggy, and keep long hair tied back. Put all tools band materials away after use. Do not allow magnets to fly together from a distance, do not play with powerful magnets, keep free magnets away from powered solenoids. Table 32: A.P.E.S. payload failure modes Potential Failure Effects of Failure Failure Prevention No power Experiment cannot be Check batteries, connections, and performed switches Data doesn't record No experimental data Ensure power is connected to the payload computer and that all connections are firmly secured Magnetic field interferes with flight computer Accelerometers No experimental data Record erroneous acceleration values Shield the flight computer from any EMF interference Calibrate and test accelerometers Solenoids Too much current goes into the solenoids Experiment cannot be performed, wires melt The wires in the solenoids get very hot Check connections, ensure over heating will not occur during testing Make sure current is only pulsed into the solenoids 71 of 94

72 Potential Failure Effects of Failure Failure Prevention Improper dress during construction Maiming, cuts, scrapes, serious injury. Do not wear open toed shoes in the build lab. Keep long hair tied back. Do not wear baggy clothing. Avionics Chips or boards are manufactured incorrectly causing equipment failures and misfires Test avionics operations, and perform a flight test Vehicle Safety To assist in the elimination of risks to all members and bystanders involved the launch vehicle risks must be identified to a reasonable degree. Relevant risks will be identified through the use of a four step method and tabulated afterwards. To mitigate potential failures of the launch vehicle potential failure modes will be developed as well as ways to prevent them from taking place. The specific hazards will also be classified according to how likely it is to occur under normal operating procedures. Risks continue to be identified for the launch vehicle and payload by exploiting a four-step risk management process. This process helps by pinpointing risks that could cause damage or harm to the environment or people. The steps are listed in Table 33. Table 33: Risk Identification and Mitigation Steps Step Name Step Definition 1. Hazard Identification The first step is to correctly identify potential hazards that could cause serious injury or death. Hazard identification will be achieved through team safety sessions and brainstorming. 2. Risk and Hazard Assessment Every hazard will undergo extensive analysis to determine how serious the issue is and the best way to approach the issue. 3. Risk Control and Elimination After the hazards are identified and assessed a method is produced to avoid the issue. 4. Reviewing Assessments As new information becomes available the assessments will be reviewed and revised as necessary. 72 of 94

73 The steps outlined above are being used to develop a set of standard operating procedures for launch vehicle construction, payload construction, ground testing, and on all launch day safety checklists. Materials Safety Data Sheets of all materials used in construction are listed in Appendix 5. Failure modes for the launch vehicle were developed to better ensure success of the entire project. Possible modes, resultant problem, and mitigation procedures are given for each failure mode. These modes will continue to evolve and expand in scope as the project progresses. The mitigation methods will be continuously incorporated into preflight checklists. The mitigation items detailed therein will be incorporated into the preflight checklist. Launch vehicle failure modes and mitigation are listed in Table 34. Table 34: Launch vehicle failure modes and mitigation Potential Failure Effects of Failure Failure Prevention Fins Launch vehicle flight path becomes unstable Test fin failure modes at connection to launch vehicle to ensure sufficient strength Structural ribs buckle on take off Launch failure, launch vehicle destroyed, possible injury from shrapnel Wear eye wear protection, test the internal structure to ensure a factor of safety against buckling Thrust retention plate Motor casing falls out Test reliability of thrust retention plate Skin zippering Launch buttons Drogue separation Internal components are exposed to flowing air currents, launch vehicle becomes unstable Launch vehicle becomes fixed to launch rail, or buttons shear off Main shoot takes full brunt of launch vehicle inertia, launch vehicle becomes ballistic Test skin adhesion reliability Ensure buttons slide easily in launch rail, ensure rail is of the proper size Do a ground test of drogue separation as well as a flight test 73 of 94

74 Potential Failure Effects of Failure Failure Prevention Main shoot Launch vehicle becomes ballistic, severe injury, irrecoverable launch vehicle Do a ground test of main shoot deployment, as well as a flight test. Land directly on fins Fins break, and launch vehicle cannot be flow twice without fixing Test fin failure modes at connection to launch vehicle to ensure sufficient strength Ignition failure Launch vehicle does not launch Follow proper procedure when setting up launch vehicle ignition system Motor failure Motor explodes Install motors properly according to NAR standards 74 of 94

75 10. Budget Our team is in contact with many Georgia Tech faculty members, many of whom are our professors, for their expertise as well as any resources they can spare. To gain additional funds needed to build the rocket and the scientific experiment along, as well as to pay for travel expenses, we plan to solicit help from local companies. For companies that donate money to our Georgia Tech USLI team, they will be acknowledged in the team reports. For those that donate large sums of money will have their logo placed on the rocket. For the rocket test launches, the team plans to work with SoAR to coordinate launches and minimize costs As shown in Table 35, the total budget for the is $5,500. Of that amount, $1,000 dollars from the Georgia Tech is specifically allocated for rocket motors. Additionally, Figure 48 illustrates the percentage breakdown of the budget while Table 36 lists the actual money allotted to each subsystem. Table 35. Sponsors of the Mile High Yellow Jackets. Organization GA Space Grant $3,500 Consortium GA Tech AE $1,000 Department GA Tech SGA $1,000 Total: $5,500 Amt. Table 36. Subsystem budget breakdown Subsystem Launch Vehicle & $2,750 Motors Flight Systems $1,750 Operations $1,000 Total: $5,500 Amt. Figure 48. Breakdown of the Mile High Yellow Jackets Budget. 75 of 94

76 11. Project Scheduling The Mile High Yellow Jacket s project is driven by the design milestone s set forth by the USLI Program Office. These milestones and their dates are listed in Table 37. It is important to note that due to the complexities of both the Mile high Yellow Jackets rocket and payload designs, the Gantt chart will contain only high-level activities. In order to visualize the major tasks/steps in our design, the Team will utilize a PERT Chart/Network Diagram. This will allow for the identification of the critical path(s) that will ensure a successful launch. Table 37. Design milestones set by the USLI Program Office. Milestone Date Proposal 26 SEP Team Selection 17 OCT Web Presence Established 4 NOV PDR Documentation 28 NOV PDR Telecon 5-14 DEC CDR Documentation 23 JAN CDR Telecon FRR Documention FRR Telecon Competition PLAR Documentation 1-10 FEB 26 MAR 2-11 APR APR 7 MAY 76 of 94

77 77 of 94

78 78 of 94

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80 12. Educational Outreach The goal of Georgia Tech s outreach program is to promote interest in the Science, Technology, Engineering, and Mathematics (STEM) fields. The intend to conduct various outreach programs targeting middle school Students and Educators. The will have an outreach request form on their webpage for Educators to request presentations or hands-on activities for their classroom. Young Astronauts Program The are planning to work in conjunction with the Georgia Tech Space Systems Design Lab (SSDL) to put on the Young Astronauts program at Madras Middle School in Newnan, Ga. The intent of this program is to expose Middle School Students to various topics in the Aerospace and STEM fields. This will be accomplished by meeting twice a month and discussing a topic followed by a related hands-on project that actively engages both Students and Educators. FIRST Lego League FIRST Lego League is an engineering competition designed for middle school children in which they build and compete with an autonomous MINDSTORMS robot. Every year there is a new competition centered on a theme exploring a real-world problem. The plan to have a booth at the Georgia State FIRST Lego League Tournament and illustrate how the skills and ideas utilized in the competition translate to real world applications, like a launch vehicle with autonomous capabilities. In addition we plan to help judge the tournament. 80 of 94

81 References Niskanen, Sampo. OpenRocket vehicle Technical Documentation. 18 July Web. Apke, Ted. "Black Powder Usage." (2009). Print. < PerfecFlite. StratoLogger SL100 Users Manual. Andover, NH: Print. < Roensch, S. (2010). "Finite Element Analysis: Introduction." 2011, from G-10 Fiberglass Epoxy Laminate Sheet. MATWEB.com. search/datasheet_print.aspx?matguid=8337b2d050d44da1b8a9a5e61b0d5f85 "Shape Effects on Drag." NASA Web. 19 Nov < 12/airplane/shaped.html>. Cavcar, Mustafa. "Compressibility Effects on Airfoil Aerodynamics." (2005). Print. "Apogee Paramagnetic Oxygen Gas Experimental Electromagnetic Separator: Preliminary Design Review." Comp. Georgia Tech University Student Launch Initiative. Atlanta: Print. Cheng, David. Field and Wave Electromagnetics. 1st ed. Reading, MA: Addison-Wesley Publishing Company, Print. 81 of 94

82 Appendix 1: Project A.P.E.S. Schedule 82 of 94

83 ID Task Name 1 Project A.P.E.S. 2 RFP Released by NASA 3 Proposal Aug '11 Sep '11 Oct '11 Nov '11 Dec '11 Jan '12 Feb '12 Mar '12 Apr '12 May '12 Jun ' Project A.P.E.S. RFP Released by NASA Proposal 4 Team Formation 5 Initial Rocket Design 6 Flight Experiment Definition 7 Internal Proposal Review 8 Proposal Submitted 9 Prelimary Design Review Proposal Submitted 9/26 Prelimary Design Review 10 Rocket 11 Structure Test Rig Design and Fabrication 12 Rocket Stucture Design and Simulations 13 Rocket Structure Component Testing 14 Rocket Structure GO/NO-GO Internal Review 15 Flight Systems Rocket Structure GO/NO-GO Internal Review 10/27 16 Component Downselection 17 Scientifc Theory Refinement 18 Experiment Modeling and Preliminary Design 19 Control System Preliminary Design 20 Preliminary Ground Testing 21 Project Level 22 Establish Web Presence 23 Fundrasing / Sponsorhip 24 PDR Documentation Submitted PDR Documentation Submitted 11/28 Project: USLI Gnatt Chart Date: Mon 9/26/11 Task Split Progress Milestone Summary Project Summary External Tasks External Milestone Deadline Page 1

84 ID Task Name 25 PDR Telecon 26 Critical Design Review Aug '11 Sep '11 Oct '11 Nov '11 Dec '11 Jan '12 Feb '12 Mar '12 Apr '12 May '12 Jun ' Critical Design Review 27 Rocket 28 Full-Scale Structure Build 29 Recovery System Ground Test 30 Competition Rocket Build 31 Initital Test Flight 32 Flight Systems 33 Flight Systems Integration 34 Control System Detailed Design 35 Detailed Experiment Modeling 36 Experiment Refinement 37 Project Level Project Level 38 Website Updates 39 Outreach Events 40 CDR Documentation Submitted 41 CDR Telecon 42 Flight Readiness Review CDR Documentation Submitted 1/23 Flight Readiness Review 43 Rocket 44 Final Test Preparation 45 Final Test Flight 46 Competition Launch Preparation 47 Flight Systems 48 Experiment Refinement Project: USLI Gnatt Chart Date: Mon 9/26/11 Task Split Progress Milestone Summary Project Summary External Tasks External Milestone Deadline Page 2

85 ID Task Name 49 Control System Refinement Aug '11 Sep '11 Oct '11 Nov '11 Dec '11 Jan '12 Feb '12 Mar '12 Apr '12 May '12 Jun ' Project Level 51 Website Updates 52 Outreach Events 53 FRR Documentation Submitted 54 FRR Telecon 55 Competition Launch 56 Post-Launch Assument Review Submitted FRR Documentation Submitted 3/26 Competition Launch 4/22 Post-Launch Assument Review Submitt 5/7 Project: USLI Gnatt Chart Date: Mon 9/26/11 Task Split Progress Milestone Summary Project Summary External Tasks External Milestone Deadline Page 3

86 Appendix 2: Launch Checklist 1. Prepare payload bay a. Ensure batteries and switches are wired to the altimeters correctly b. Ensure batteries, power supply, switch, data recorder and pressure sensors are wired correctly c. Install fresh batteries into battery holders and secure with tape d. Insert altimeter and payload into the payload bay e. Connect appropriate wires f. Verify payload powers on correctly and is working properly. If it is not, check all wires and connections g. Turn off payload power h. Arm altimeters with output shorted to verify jumper settings. This is to check battery voltage and continuity i. Disarm altimeter, un-short outputs j. Close altimeter bay 2. Assemble charges a. Test e-match resistance and make sure it is within spec b. Remove protective cover from e-match c. Measure amount of black powder determined in testing d. Put e-match on tape with sticky side up e. Pour black powder over e-match f. Seal tape g. Retest e-match 3. Ensure altimeter is disarmed 4. Connect charges to altimeter bay 5. Turn on altimeter and verify continuity 6. Disarm altimeter 7. Connect drogue shock cord to booster section and altimeter bay 8. Fold excess shock cord so it does not tangle 9. Add cellulose wadding to ensure only the Kevlar shock chord is exposed to ejection charge 10. Insert altimeter bay into drogue section and secure with shear pins 11. Pack main chute 12. Attach main shock cord to payload bay 13. Fold excess shock cord so it does not tangle 14. Add wadding under main chute and shock cord ensuring that only the Kevlar part of the shock cord will be exposed to the ejection charge 15. Attach altimeter bay to the main section with nylon rivets 16. Connect shock cord to nose cone, install nose cone and secure with shear pins 86 of 94

87 17. Assemble motor a. Follow manufacturer's instructions b. Do not get grease on propellant or delay 18. Do not install igniter until at pad 19. Install motor in launch vehicle 20. Secure positive motor retention 21. Inspect launch vehicle. Check CG to make sure it is in safe range; add nose weight if necessary 22. Arm altimeter and ensure both charges read continuity 23. Disarm altimeter 24. Bring launch vehicle to the range safety officer (RSO) table for inspection 25. Bring launch vehicle to pad, install on pad, verify that it can move freely (use a standoff if necessary) 26. Install igniter in launch vehicle 27. Touch igniter clips together to make sure they will not fire igniter when connected 28. Make sure clips are not shorted to each other or blast deflector 29. Arm altimeters via switches and wait for continuity check for both 30. Turn on payload via a switch and start stopwatches 31. Return to front line 32. Launch. Stop the stopwatches and record time from arming payload and launch 33. Watch flight so launch vehicle does not get lost 34. Recover launch vehicle 35. Disarm altimeter(s) if there are unfired charges 36. Disassemble launch vehicle, clean motor case, other parts, inspect for damage 37. Record altimeter data 38. Download payload data 87 of 94

88 Appendix 3: Ground Test Plan Ground Test Plan Goals The A.P.E.S. ground test data will provide the basis for empirical modeling of the force interactions for various configurations of the experiment at various voltages. All actions will be incremented to allow for a detailed model for extrapolation and interpolation of the data for future flight control systems. Goals are detailed in Table 38: Ground Test goals. Table 38: Ground Test goals Ground Test Goal Ground Test Goal Definition 1 Calibrate Sensing Data 2 Map Magnetic Fields 3 Detect force equilibrium 4 Develop model for control of voltage Test Sequence 1 Calibration The static field of a solenoid in the ground test platform configuration to be tested will be mapped at RMS voltage increments of 0.25 V from 0.25 V to 6 V. For each RMS Voltage, no fewer than nine (9) points of data shall be produced with Hall-effect sensors, measuring 3D field vectors on at least three (3) sets of axes. The IR distance sensor will be calibrated. Materials of varying IR reflectivity will be used and the results will be compared to provide a recommendation for the marking of the test article. The Camera Cube will be calibrated with visual targets. Test Sequence 2 1-Dimensional testing Equilibrium testing with no internal magnetism. A single vertically-oriented solenoid will be utilized to lift the test article covered in IR reflective material to equilibrium points within a cylinder, from 1 cm to 7 cm in steps of 1 cm. Hall-effect sensors will be used to map fields at 88 of 94

89 each equilibrium point identically to the static field mapping. The IR sensor will detect distance from below the cylinder. The IR sensor will be lowered to the minimum read distance using MakerBeam. The minimum read distance shall be confirmed by data sheets and calibration. Test Sequence 3 Equilibrium de-scope testing with internal magnetism. One (1) neodymium magnet shall be placed in the center of the test article and covered with IR reflective material. A single vertically-oriented solenoid will be utilized to lift the test article to equilibrium points within a cylinder, from 1 cm to 7 cm in steps of 1 cm. Hall-effect sensors will be used to map fields at each equilibrium point identically to the static field mapping. The test setup should allow for both pulling of the test article as well as pushing of the test article. Test Sequence 4 Similar testing will be completed using a horizontal sheet with the sample placed on top. Solenoids will be used to pull and hold the sample in the middle of the platform at equilibrium. These tests will be completed with the permanent magnet sample. The Camera Cube will allow for object detection. Fields will be mapped at equilibrium. Test Sequence 5 A permanently magnetic test article will be levitated from rest in 3-dimensions to equilibrium at central points in the test stand. Incrementing of the equilibrium point will allow for greater control of the test article. Object detection will be accomplished with the Camera Cube. Fields will be mapped at equilibrium. Test Sequence 6 The flight model will be tested and disturbances will be introduced. 89 of 94

90 Appendix 4: Derivation of Force Equations for a Ferrite Platform Derivation of Force Equation for Ferrite Platform The force due to the field of a magnetic dipole of moment m that is in a material of permeability is on a magnetic dipole of moment ( ) [( ) ( ) ( ) ( )( ) ] where r is the vector from to m, and is again the unit vector in the direction of r. (14) At any location inside the platform, the microscopic magnetic dipole moment is. Therefore, the total force on the platform is given by the following equation ( ) [( ) ( ) ( ) ( )( ) ] As it is assumed that the ferrite has constant susceptibility, (15) ( ( ) ) (16) where the field is generated by a solenoid of Equation (2), it can be found that. Substituting this equation in for M in ( ) [( )( ( ) ) (( ( ) ) ) ( ( ( ) )) (17) ( )(( ( ) ) ) ] ( ) [( ( ) ( ) ) ( ( ) ( ) ) ( ( ) ( ) ) (18) ( )( ( ) ( )) ] ( ) [( ( ) ( ) ) And therefore, ( ( ) ) ( ( ) ( ) ) ( ) ] (19) 90 of 94

91 ( ) [( ) ( ) ( ) ] which was the result that was to be derived. (20) 91 of 94

92 Appendix 5: Flight Computer Schematic 92 of 94

93 HMC1043 ADXL345 ATMEGA2560AU XBEE-1B1 GND GND GND GND GND GND AVR_SPI_PRG_6PTH GND VCC VCC VCC GND VCC GND GND VCC VCC GND MC33269ST-3.3T3 MC33269D-5.0 M7 100n 47u 47u GND GND GND GND VCC VCC VCCVCC GND GND 27R 1M 100n 100n 100n 100n 1K HLMP6 GND VCC GND VCC 22p VCC LM358D LM358D 10K 10K 10K 10K VCC GND 100n GND GND 100n VCC GND 100n GND P9 P9 P10 P10 P11 P11 P12 P12 P13 P13 P14 P14 P15 P15 P16 P16 P1 P1 P2 P2 P3 P3 P4 P4 P5 P5 P6 P6 P7 P7 P8 P8 U$1 VDD 1 NC 11 NC 3 GND 4 GND 5 VSS 6 CS 7 INT1 8 INT2 9 GND 2 NC 10 SDO 12 SDA 13 SCL 14 U1 (A8)PC0 53 (A9)PC1 54 (A10)PC2 55 (A11)PC3 56 (A12)PC4 57 (A13)PC5 58 (A14)PC6 59 (A15)PC7 60 (AD0)PA0 78 (AD1)PA1 77 (AD2)PA2 76 (AD3)PA3 75 (AD4)PA4 74 (AD5)PA5 73 (AD6)PA6 72 (AD7)PA7 71 (ADC0)PF0 97 (ADC1)PF1 96 (ADC2)PF2 95 (ADC3)PF3 94 (ADC4/TCK)PF4 93 (ADC5/TMS)PF5 92 (ADC6/TDO)PF6 91 (ADC7/TDI)PF7 90 (ALE)PG2 70 (CLKO/ICP3/INT7)PE7 9 (ICP1)PD4 47 (MISO/PCINT3)PB3 22 (MOSI/PCINT2)PB2 21 (OC0A/OC1C/PCINT7)PB7 26 (OC0B)PG5 1 (OC1A/PCINT5)PB5 24 (OC1B/PCINT6)PB6 25 (OC2A/PCINT4)PB4 23 (OC3A/AIN1)PE3 5 (OC3B/INT4)PE4 6 (OC3C/INT5)PE5 7 (RD)PG1 52 (RXD0/PCIN8)PE0 2 (RXD1/INT2)PD2 45 (SCK/PCINT1)PB1 20 (SCL/INT0)PD0 43 (SDA/INT1)PD1 44 (SS/PCINT0)PB0 19 (T0)PD7 50 (T1)PD6 49 (T3/INT6)PE6 8 (TOSC1)PG4 29 (TOSC2)PG3 28 (TXD0)PE1 3 (TXD1/INT3)PD3 46 (WR)PG0 51 (XCK0/AIN0)PE2 4 (XCK1)PD5 48 AGND 99 AREF 98 AVCC 100 GND PH0(RXD2) 12 PH1(TXD2) 13 PH2(XCK2) 14 PH3(OC4A) 15 PH4(OC4B) 16 PH5(OC4C) 17 PH6(OC2B) 18 PH7(T4) 27 PJ0(RXD3/PCINT9) 63 PJ1(TXD3/PCINT10) 64 PJ2(XCK3/PCINT11) 65 PJ3(PCINT12) 66 PJ4(PCINT13) 67 PJ5(PCINT14) 68 PJ6(PCINT15) 69 PJ7 79 PK0(ADC8/PCINT16) 89 PK1(ADC9/PCINT17) 88 PK2(ADC10/PCINT18) 87 PK3(ADC11/PCINT19) 86 PK4(ADC12/PCINT20) 85 PK5(ADC13/PCINT21) 84 PK6(ADC14/PCINT22) 83 PK7(ADC15/PCINT23) 82 PL0(ICP4) 35 PL1(ICP5) 36 PL2(T5) 37 PL3(OC5A) 38 PL4(OC5B) 39 PL5(OC5C) 40 PL6 41 PL7 42 RESET 30 VCC XTAL1 34 XTAL2 33 U2 GND GND GND VCC VCC VCC VDD 1 DOUT 2 DIN 3 DIO12 4 RESET 5 RSSI 6 DIO11 7 RES 8 DTR 9 GND 10 DIO4 11 CTS 12 DIO9 13 RES 14 DIO5 15 RTS 16 DIO3 17 DIO2 18 DIO1 19 DIO0 20 U$3 LED1 R1 R2 R3 R J1 Fastrax UP501 GPS Receiver RXD P$1 TXD P$2 GND P$3 VDD P$4 VDD_B P$5 PPS P$6 Q1 C5 C S POWER UP IN 3 OUT 4 2 IC2 VI 3 1 VO 2 IC3 ADJ 1 2 JP1 D1 C7 C8 C JP2 R7 R8 C10 C11 C12 C13 R5 BLUEON C IC1A IC1B RN1A 2 7 RN1B 3 6 RN1C 4 5 RN1D 1 8 RN2A 2 7 RN2B 3 6 RN2C 4 5 RN2D C3 C4 T1 IN 1 EN 3 NC/FB 4 OUT 5 GND 2 C14 MISO MISO MISO SCK SCK SCK RESET RESET RESET RESET RESET MOSI MOSI MOSI AREF AREF XTAL2 XTAL2 XTAL1 XTAL1 VIN VIN +3V3 +3V3 +3V3 +3V3 +3V3 +3V3 +3V3 +5V +5V +5V +5V +5V +5V +5V SDA SDA SCL SCL USBVCC MOSI RESET SCK MISO +5 GND + +

94 Appendix 6: Material Safety Data Sheets (MSDS) 94 of 94

95 Material Safety Data Sheet Ferrosoferric Oxide, Black Powder MSDS Section 1: Chemical Product and Company Identification Product Name: Ferrosoferric Oxide, Black Powder Catalog Codes: SLF1477 CAS#: RTECS: Not available. TSCA: TSCA 8(b) inventory: Ferrosoferric Oxide, Black Powder CI#: Not available. Contact Information: Sciencelab.com, Inc Smith Rd. Houston, Texas US Sales: International Sales: Order Online: ScienceLab.com CHEMTREC (24HR Emergency Telephone), call: Synonym: Iron Oxide International CHEMTREC, call: Chemical Name: Ferrosoferric Oxide, Black Powder For non-emergency assistance, call: Chemical Formula: Fe3-O4 Section 2: Composition and Information on Ingredients Composition: Name CAS # % by Weight Ferrosoferric Oxide, Black Powder Toxicological Data on Ingredients: Ferrosoferric Oxide, Black Powder: ORAL (LD50): Acute: 5000 mg/kg [Rat]. Section 3: Hazards Identification Potential Acute Health Effects: Slightly hazardous in case of skin contact (irritant), of eye contact (irritant),. Non-irritant for lungs. Potential Chronic Health Effects: CARCINOGENIC EFFECTS: Classified None. by NTP, None. by OSHA, None. by NIOSH. MUTAGENIC EFFECTS: Not available. TERATOGENIC EFFECTS: Not available. DEVELOPMENTAL TOXICITY: Not available. The substance is toxic to lungs, upper respiratory tract. Repeated or prolonged exposure to the substance can produce target organs damage. Section 4: First Aid Measures p. 1

96 Eye Contact: No known effect on eye contact, rinse with water for a few minutes. Skin Contact: After contact with skin, wash immediately with plenty of water. Gently and thoroughly wash the contaminated skin with running water and non-abrasive soap. Be particularly careful to clean folds, crevices, creases and groin. Cover the irritated skin with an emollient. If irritation persists, seek medical attention. Serious Skin Contact: Not available. Inhalation: Allow the victim to rest in a well ventilated area. Seek immediate medical attention. Serious Inhalation: Not available. Ingestion: Do not induce vomiting. Loosen tight clothing such as a collar, tie, belt or waistband. If the victim is not breathing, perform mouth-to-mouth resuscitation. Seek immediate medical attention. Serious Ingestion: Not available. Flammability of the Product: Non-flammable. Auto-Ignition Temperature: Not applicable. Flash Points: Not applicable. Flammable Limits: Not applicable. Products of Combustion: Not available. Section 5: Fire and Explosion Data Fire Hazards in Presence of Various Substances: Not applicable. Explosion Hazards in Presence of Various Substances: Risks of explosion of the product in presence of mechanical impact: Not available. Risks of explosion of the product in presence of static discharge: Not available. Fire Fighting Media and Instructions: Not applicable. Special Remarks on Fire Hazards: Material is not combustible. Use extinguishing media suitable for other combustible material in the area Special Remarks on Explosion Hazards: Not available. Section 6: Accidental Release Measures Small Spill: Use appropriate tools to put the spilled solid in a convenient waste disposal container. Finish cleaning by spreading water on the contaminated surface and dispose of according to local and regional authority requirements. Large Spill: Use a shovel to put the material into a convenient waste disposal container. Finish cleaning by spreading water on the contaminated surface and allow to evacuate through the sanitary system. Precautions: Section 7: Handling and Storage p. 2

97 Do not ingest. Do not breathe dust. If ingested, seek medical advice immediately and show the container or the label. Storage: No specific storage is required. Use shelves or cabinets sturdy enough to bear the weight of the chemicals. Be sure that it is not necessary to strain to reach materials, and that shelves are not overloaded. Section 8: Exposure Controls/Personal Protection Engineering Controls: Use process enclosures, local exhaust ventilation, or other engineering controls to keep airborne levels below recommended exposure limits. If user operations generate dust, fume or mist, use ventilation to keep exposure to airborne contaminants below the exposure limit. Personal Protection: Safety glasses. Lab coat. Dust respirator. Be sure to use an approved/certified respirator or equivalent. Gloves. Personal Protection in Case of a Large Spill: Splash goggles. Full suit. Dust respirator. Boots. Gloves. A self contained breathing apparatus should be used to avoid inhalation of the product. Suggested protective clothing might not be sufficient; consult a specialist BEFORE handling this product. Exposure Limits: Not available. Section 9: Physical and Chemical Properties Physical state and appearance: Solid. (Solid powder.) Odor: Odorless. Taste: Not available. Molecular Weight: Not available. Color: Black ph (1% soln/water): Not available. Boiling Point: 1000 C (1832 F) Melting Point: Not available. Critical Temperature: Not available. Specific Gravity: 4.6 (Water = 1) Vapor Pressure: Not applicable. Vapor Density: Not available. Volatility: Not available. Odor Threshold: Not available. Water/Oil Dist. Coeff.: Not available. Ionicity (in Water): Not available. Dispersion Properties: Not available. p. 3

98 Solubility: Not available. Section 10: Stability and Reactivity Data Stability: The product is stable. Instability Temperature: Not available. Conditions of Instability: Not available. Incompatibility with various substances: Not available. Corrosivity: Not available. Special Remarks on Reactivity: Not available. Special Remarks on Corrosivity: Not available. Polymerization: No. Section 11: Toxicological Information Routes of Entry: Absorbed through skin. Dermal contact. Eye contact. Inhalation. Toxicity to Animals: Acute oral toxicity (LD50): 5000 mg/kg [Rat]. Chronic Effects on Humans: CARCINOGENIC EFFECTS: Classified None. by NTP, None. by OSHA, None. by NIOSH. The substance is toxic to lungs, upper respiratory tract. Other Toxic Effects on Humans: Slightly hazardous in case of skin contact (irritant),. Non-irritant for lungs. Special Remarks on Toxicity to Animals: Not available. Special Remarks on Chronic Effects on Humans: Not available. Special Remarks on other Toxic Effects on Humans: Not available. Ecotoxicity: Not available. BOD5 and COD: Not available. Section 12: Ecological Information Products of Biodegradation: Possibly hazardous short term degradation products are not likely. However, long term degradation products may arise. Toxicity of the Products of Biodegradation: The product itself and its products of degradation are not toxic. Special Remarks on the Products of Biodegradation: Not available. Waste Disposal: Section 13: Disposal Considerations p. 4

99 Section 14: Transport Information DOT Classification: Not a DOT controlled material (United States). Identification: Not applicable. Special Provisions for Transport: Not applicable. Section 15: Other Regulatory Information Federal and State Regulations: California prop. 65: This product contains the following ingredients for which the State of California has found to cause cancer, birth defects or other reproductive harm, which would require a warning under the statute: Ferrosoferric Oxide, Black Powder Massachusetts RTK: Ferrosoferric Oxide, Black Powder New Jersey: Ferrosoferric Oxide, Black Powder TSCA 8(b) inventory: Ferrosoferric Oxide, Black Powder Other Regulations: OSHA: Hazardous by definition of Hazard Communication Standard (29 CFR ). EINECS: This product is on the European Inventory of Existing Commercial Chemical Substances. Other Classifications: WHMIS (Canada): CLASS D-2B: Material causing other toxic effects (TOXIC). DSCL (EEC): This product is not classified according to the EU regulations. HMIS (U.S.A.): Health Hazard: 1 Fire Hazard: 0 Reactivity: 0 Personal Protection: E National Fire Protection Association (U.S.A.): Health: 1 Flammability: 0 Reactivity: 0 Specific hazard: Protective Equipment: Gloves. Lab coat. Dust respirator. Be sure to use an approved/certified respirator or equivalent. Wear appropriate respirator when ventilation is inadequate. Safety glasses. Section 16: Other Information p. 5

100 References: Not available. Other Special Considerations: Not available. Created: 10/09/ :33 PM Last Updated: 11/06/ :00 PM The information above is believed to be accurate and represents the best information currently available to us. However, we make no warranty of merchantability or any other warranty, express or implied, with respect to such information, and we assume no liability resulting from its use. Users should make their own investigations to determine the suitability of the information for their particular purposes. In no event shall ScienceLab.com be liable for any claims, losses, or damages of any third party or for lost profits or any special, indirect, incidental, consequential or exemplary damages, howsoever arising, even if ScienceLab.com has been advised of the possibility of such damages. p. 6

101 Material Safety Data Sheet Ammonium perchlorate MSDS Section 1: Chemical Product and Company Identification Product Name: Ammonium perchlorate Catalog Codes: SLA2725 CAS#: RTECS: SC TSCA: TSCA 8(b) inventory: Ammonium perchlorate CI#: Not available. Synonym: Chemical Formula: NH4ClO4 Contact Information: Sciencelab.com, Inc Smith Rd. Houston, Texas US Sales: International Sales: Order Online: ScienceLab.com CHEMTREC (24HR Emergency Telephone), call: International CHEMTREC, call: For non-emergency assistance, call: Section 2: Composition and Information on Ingredients Composition: Name CAS # % by Weight Ammonium perchlorate Toxicological Data on Ingredients: Ammonium perchlorate LD50: Not available. LC50: Not available. Section 3: Hazards Identification Potential Acute Health Effects: Hazardous in case of skin contact (irritant), of eye contact (irritant), of ingestion, of inhalation. Prolonged exposure may result in skin burns and ulcerations. Over-exposure by inhalation may cause respiratory irritation. Potential Chronic Health Effects: Hazardous in case of skin contact (irritant), of eye contact (irritant), of ingestion, of inhalation. CARCINOGENIC EFFECTS: Not available. MUTAGENIC EFFECTS: Not available. TERATOGENIC EFFECTS: Not available. DEVELOPMENTAL TOXICITY: Not available. The substance is toxic to blood, kidneys. Repeated or prolonged exposure to the substance can produce target organs damage. Section 4: First Aid Measures p. 1

102 Eye Contact: Check for and remove any contact lenses. In case of contact, immediately flush eyes with plenty of water for at least 15 minutes. Cold water may be used. Get medical attention. Skin Contact: In case of contact, immediately flush skin with plenty of water. Cover the irritated skin with an emollient. Remove contaminated clothing and shoes. Cold water may be used.wash clothing before reuse. Thoroughly clean shoes before reuse. Get medical attention. Serious Skin Contact: Wash with a disinfectant soap and cover the contaminated skin with an anti-bacterial cream. Seek medical attention. Inhalation: If inhaled, remove to fresh air. If not breathing, give artificial respiration. If breathing is difficult, give oxygen. Get medical attention. Serious Inhalation: Evacuate the victim to a safe area as soon as possible. Loosen tight clothing such as a collar, tie, belt or waistband. If breathing is difficult, administer oxygen. If the victim is not breathing, perform mouth-to-mouth resuscitation. Seek medical attention. Ingestion: Do NOT induce vomiting unless directed to do so by medical personnel. Never give anything by mouth to an unconscious person. If large quantities of this material are swallowed, call a physician immediately. Loosen tight clothing such as a collar, tie, belt or waistband. Serious Ingestion: Not available. Section 5: Fire and Explosion Data Flammability of the Product: May be combustible at high temperature. Auto-Ignition Temperature: Not available. Flash Points: Not available. Flammable Limits: Not available. Products of Combustion: Not available. Fire Hazards in Presence of Various Substances: Flammable in presence of shocks, of heat, of reducing materials, of combustible materials, of organic materials. Explosion Hazards in Presence of Various Substances: Extremely explosive in presence of open flames and sparks, of shocks, of heat, of reducing materials, of organic materials. Slightly explosive in presence of acids. Fire Fighting Media and Instructions: Oxidizing material. Do not use water jet. Use flooding quantities of water. Avoid contact with organic materials. Special Remarks on Fire Hazards: Not available. Special Remarks on Explosion Hazards: Not available. Section 6: Accidental Release Measures p. 2

103 Small Spill: Use appropriate tools to put the spilled solid in a convenient waste disposal container. Large Spill: Oxidizing material. Stop leak if without risk. Avoid contact with a combustible material (wood, paper, oil, clothing...). Keep substance damp using water spray. Do not touch spilled material. Prevent entry into sewers, basements or confined areas; dike if needed. Eliminate all ignition sources. Call for assistance on disposal. Section 7: Handling and Storage Precautions: Keep away from heat. Keep away from sources of ignition. Keep away from combustible material.. Empty containers pose a fire risk, evaporate the residue under a fume hood. Ground all equipment containing material. Do not breathe dust. Take precautionary measures against electrostatic discharges. Wear suitable protective clothing. In case of insufficient ventilation, wear suitable respiratory equipment. If you feel unwell, seek medical attention and show the label when possible. Avoid contact with skin and eyes. Keep away from incompatibles such as reducing agents, combustible materials, organic materials, acids. Storage: Keep container tightly closed. Keep container in a cool, well-ventilated area. Separate from acids, alkalies, reducing agents and combustibles. See NFPA 43A, Code for the Storage of Liquid and Solid Oxidizers. Section 8: Exposure Controls/Personal Protection Engineering Controls: Use process enclosures, local exhaust ventilation, or other engineering controls to keep airborne levels below recommended exposure limits. If user operations generate dust, fume or mist, use ventilation to keep exposure to airborne contaminants below the exposure limit. Personal Protection: Splash goggles. Lab coat. Dust respirator. Be sure to use an approved/certified respirator or equivalent. Gloves. Personal Protection in Case of a Large Spill: Splash goggles. Full suit. Dust respirator. Boots. Gloves. A self contained breathing apparatus should be used to avoid inhalation of the product. Suggested protective clothing might not be sufficient; consult a specialist BEFORE handling this product. Exposure Limits: Not available. Section 9: Physical and Chemical Properties Physical state and appearance: Solid. (Crystals solid.) Odor: Not available. Taste: Not available. Molecular Weight: g/mole Color: Colorless. ph (1% soln/water): Not available. Boiling Point: Not available. Melting Point: Decomposes. p. 3

104 Critical Temperature: Not available. Specific Gravity: 1.95 (Water = 1) Vapor Pressure: Not applicable. Vapor Density: Not available. Volatility: Not available. Odor Threshold: Not available. Water/Oil Dist. Coeff.: Not available. Ionicity (in Water): Not available. Dispersion Properties: See solubility in water, methanol, acetone. Solubility: Soluble in cold water, methanol. Partially soluble in acetone. Insoluble in diethyl ether. Stability: Unstable. Instability Temperature: Not available. Conditions of Instability: No additional remark. Section 10: Stability and Reactivity Data Incompatibility with various substances: Extremely reactive or incompatible with reducing agents, combustible materials, organic materials. Reactive with acids. Corrosivity: Non-corrosive in presence of glass. Special Remarks on Reactivity: Not available. Special Remarks on Corrosivity: Not available. Polymerization: Will not occur. Routes of Entry: Eye contact. Inhalation. Ingestion. Toxicity to Animals: LD50: Not available. LC50: Not available. Section 11: Toxicological Information Chronic Effects on Humans: Causes damage to the following organs: blood, kidneys. Other Toxic Effects on Humans: Hazardous in case of skin contact (irritant), of ingestion, of inhalation. Special Remarks on Toxicity to Animals: Not available. Special Remarks on Chronic Effects on Humans: Not available. Special Remarks on other Toxic Effects on Humans: Not available. p. 4

105 Section 12: Ecological Information Ecotoxicity: Not available. BOD5 and COD: Not available. Products of Biodegradation: Possibly hazardous short/long term degradation products are to be expected. Toxicity of the Products of Biodegradation: The products of degradation are more toxic. Special Remarks on the Products of Biodegradation: Not available. Waste Disposal: Section 13: Disposal Considerations Section 14: Transport Information DOT Classification: CLASS 5.1: Oxidizing material. Identification: : Ammonium Perchlorate UNNA: UN1442 PG: II Special Provisions for Transport: Not available. Federal and State Regulations: Pennsylvania RTK: Ammonium perchlorate Massachusetts RTK: Ammonium perchlorate TSCA 8(b) inventory: Ammonium perchlorate Section 15: Other Regulatory Information Other Regulations: OSHA: Hazardous by definition of Hazard Communication Standard (29 CFR ). Other Classifications: WHMIS (Canada): CLASS C: Oxidizing material. DSCL (EEC): R8- Contact with combustible material may cause fire. R36/38- Irritating to eyes and skin. HMIS (U.S.A.): Health Hazard: 2 Fire Hazard: 1 Reactivity: 4 Personal Protection: E National Fire Protection Association (U.S.A.): Health: 2 Flammability: 1 Reactivity: 4 p. 5

106 Specific hazard: Protective Equipment: Gloves. Lab coat. Dust respirator. Be sure to use an approved/certified respirator or equivalent. Wear appropriate respirator when ventilation is inadequate. Splash goggles. Section 16: Other Information References: Not available. Other Special Considerations: Not available. Created: 10/09/ :57 PM Last Updated: 11/06/ :00 PM The information above is believed to be accurate and represents the best information currently available to us. However, we make no warranty of merchantability or any other warranty, express or implied, with respect to such information, and we assume no liability resulting from its use. Users should make their own investigations to determine the suitability of the information for their particular purposes. In no event shall ScienceLab.com be liable for any claims, losses, or damages of any third party or for lost profits or any special, indirect, incidental, consequential or exemplary damages, howsoever arising, even if ScienceLab.com has been advised of the possibility of such damages. p. 6

107 MATERIAL SAFETY DATA SHEET Prepared to U.S. OSHA, CMA, ANSI and Canadian WHMIS Standards.This Material Safety Data Sheet is offered pursuant to OSHA s Hazard Communication Standard (29 CFR ). Other government regulations must be reviewed for applicability to these products. WARNING: PRODUCT COMPONENTS PRESENT HEALTH AND SAFETY HAZARDS. READ AND UNDERSTAND THIS MATERIAL SAFETY DATA SHEET (M.S.DS.). ALSO, FOLLOW YOUR EMPLOYER S SAFETY PRACTICES. This product may contain Chromium and/or Nickel which are listed by OSHA, NTP, or IARC as being a carcinogen or potential carcinogen. Use of this product may expose you or others to fumes and gases at levels exceeding those established by the American Conference of Governmental Industrial Hygienists (ACGIH) or the Occupational Safety and Health Administration (OSHA) The information contained herein relates only to the specific product. If the product is combined with other materials, all component properties must be considered. BE SURE TO CONSULT THE LATEST VERSION OF THE MSDS. MATERIAL SAFETY DATA SHEETS ARE AVAILABLE FROM HARRIS PRODUCTS GROUP salesinfo@jwharris.com STATEMENT OF LIABILITY-DISCLAIMER To the best of the Harris Products Group knowledge, the information and recommendations contained in this publication are reliable and accurate as of the date prepared. However, accuracy, suitability, or completeness are not guaranteed, and no warranty, guarantee, or representation, expressed or implied, is made by Harris Products Group as to the absolute correctness or sufficiency of any representation contained in this and other publications; Harris Products Group. assumes no responsibility in connection therewith; nor can it be assumed that all acceptable safety measures are contained in this and other publications, or that other or additional measures may not be required under particular or exceptional conditions or circumstances. Data may be changed from time to time. PART I What is the material and what do I need to know in an emergency? 1. PRODUCT IDENTIFICATION TRADE NAME (AS LABELED): LEAD SOLDER CHEMICAL NAME/CLASS: Metal Alloy SYNONYMS: Not Applicable PRODUCT USE: Soldering DOCUMENT NUMBER: 0126 SUPPLIER/MANUFACTURER'S NAME: HARRIS PRODUCTS GROUP ADDRESS: 4501 Quality Place, Mason, Ohio EMERGENCY PHONE: CHEMTREC: BUSINESS PHONE: FAX DATE OF PREPARATION: July 12, 2007 REVIEWED October 22, COMPOSITION and INFORMATION ON INGREDIENTS NOMINAL COMPOSITION WEIGHT % Wire TRADE NAME 30/70 40/60 50/50 60/40 63/37 70/30 90/10 Tin (Sn) 30% 40% 50% 60% 63% 70% 90% Lead (Pb) 70% 60% 50% 40% 37% 30% 10% NOMINAL COMPOSITION WEIGHT % Flux Core TRADE NAME Activated Rosin CAS # Ammonium Chloride CAS # Zinc Chloride CAS # Water CAS # ACID CORE < 20% < 70% Balance ROSIN CORE 100% LEAD SOLDERS October 22, 2004 PAGE 1 OF 13